Aerodynamic heat and thermal response coupling calculation method considering ablation pyrolysis product injection

Abstract

The invention discloses an aerodynamic heat and thermal response coupling calculation method considering ablation pyrolysis product injection, relates to the field of aerodynamic heat and thermal protection coupling numerical calculation, and establishes an aerodynamic heat environment numerical calculation method under a mass injection condition. A thermal response numerical calculation method considering a surface ablation reaction, an internal pyrolysis reaction, pyrolysis gas flow and solid heat conduction; the invention relates to an aerodynamic heat / thermal response coupling calculation method considering ablation pyrolysis product injection, and a coupling data transmission method strategy between the two. According to the method, the generation, transportation and injection processes of surface ablation reaction products and internal pyrolysis reaction products of the high-speed aircraft ablation pyrolysis thermal protection system can be simulated, and a coupled aerodynamic thermal / thermal response result of the high-speed aircraft ablation pyrolysis thermal protection system considering mass injection can be calculated.

Description

Technical Field

[0001] This invention relates to the field of coupled numerical calculation of aerodynamic heat and thermal protection, and more specifically, to a method for coupled calculation of aerodynamic heat and thermal response considering the ejection of ablation pyrolysis products. Background Technology

[0002] This invention belongs to the fields of aerodynamic thermal and thermal protection technology and fluid-thermal-solid coupled computation. It can couple and simulate the aerodynamic thermal environment and material thermal response under surface ablation and internal pyrolysis gas ejection conditions of thermal protection systems for high-speed (Mach number greater than 5) aircraft. When an aircraft flies at high speed within the atmosphere, it faces a severe aerodynamic thermal environment. To ensure the thermal safety of the aircraft, ablation thermal protection systems are often used. Ablation thermal protection systems have been around for a long time and are now widely used in various high-speed (Mach number greater than 5) aircraft, such as ballistic missiles, reentry capsules / entry vehicles, and near-space vehicles. The working principle of an ablation thermal protection system is that materials undergo ablation pyrolysis under harsh aerodynamic thermal environments, resulting in mass loss of the material, absorbing and carrying away a large amount of heat, and preventing some heat transfer into the aircraft. The internal pyrolysis, surface ablation reaction, melting, evaporation / sublimation processes in the ablation thermal protection system all involve the surface mass ejection phenomenon of fluid leaving the material surface and entering the flow field. There is an interaction between the aerodynamic thermal environment and the structural thermal response of high-speed (Mach number greater than 5) aircraft. In particular, when considering the fluid domain of chemical non-equilibrium flow, the solid domain of ablation pyrolysis materials, and the mass ejection behavior at the gas-solid interface, the aerodynamic thermal / thermal response coupling effect becomes more compact and complex.

[0003] Existing simulation methods for aerodynamic / thermal responses considering the ejection of ablation pyrolysis products have certain limitations. In flow field simulation, some methods ignore the impact of mass ejection on the thermal environment, only setting temperature boundary conditions; others, to simplify calculations, use a mass ejection coefficient to correct for aerodynamic environments that do not consider mass ejection. These methods do not calculate the behavior of fluid entering a high-speed flow field from the surface. While this reduces computational costs, it makes it difficult to simulate the comprehensive impact of mass ejection on the aerodynamic environment. Furthermore, when an aircraft is in a high-enthalpy inflow, chemical reactions occur within the boundary layer. Mass-ejected gases entering the boundary layer may further react chemically with these gases, thus affecting the aerodynamic environment. Currently, there are few methods that simulate this phenomenon.

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

[0005] Existing methods have resulted in a lack of effective assessment tools for cutting-edge engineering problems such as accurate prediction of aircraft thermal environment and assessment and design of thermal protection capabilities, making it difficult to guarantee the thermal safety of new high-speed (Mach number greater than 5) aircraft. Summary of the Invention

[0006] The purpose of this invention is to propose a currently lacking aerodynamic / thermal response coupling calculation method that considers the ejection of ablation pyrolysis products, providing support for the design and evaluation of thermal protection systems for high-speed (Mach number greater than 5) aircraft.

[0007] To achieve the above-mentioned objective, this invention provides a method for coupled calculation of aerodynamic heat and thermal response considering the ejection of ablation pyrolysis products, the method comprising:

[0008] Step 1: Determine the geometric shape of the aircraft and draw the flow field mesh and structural field mesh;

[0009] Step 2: Select the flight trajectory point. Based on the flight Mach number, altitude, angle of attack and sideslip angle at the corresponding moment of the selected flight trajectory point, solve the aerodynamic thermal environment based on the flow field grid to obtain the aerodynamic thermal environment calculation results.

[0010] Step 3: Interact the flow field data and structural field data at the wall location, and transfer the wall heat flux distribution and pressure distribution in the aerodynamic thermal environment calculation results to the grid points on the structural field surface to obtain the boundary conditions for thermal response calculation;

[0011] Step 4: Establish mass conservation equations considering the ablation reaction on the aircraft surface, internal pyrolysis, and pyrolysis gas flow; establish energy conservation equations considering the heat absorption of the ablation reaction on the aircraft material surface and the internal pyrolysis reaction, the heat conduction of the solid phase, and the convective heat transfer of the pyrolysis gas. Based on the boundary conditions for thermal response calculation obtained in Step 3, perform thermal response calculations using the mass conservation equations and energy conservation equations, simulate for a set duration, and obtain the temperature distribution of the aircraft structure, the ablation reaction on the aircraft surface, and the internal pyrolysis reaction and pyrolysis gas flow.

[0012] Step 5: Interact the flow field data and structural field data at the wall location. Obtain the surface temperature distribution of the aircraft based on the structural temperature distribution of the aircraft. Obtain the composition of the ablation reaction products and the ejector mass flow rate distribution of the aircraft surface based on the ablation reaction of the aircraft surface. Obtain the composition of the internal pyrolysis reaction products and the ejector mass flow rate distribution of the aircraft surface based on the internal pyrolysis reaction and pyrolysis gas flow of the aircraft. Transfer the surface temperature distribution, the composition of the ablation reaction products and the ejector mass flow rate distribution of the aircraft surface, and the composition of the internal pyrolysis reaction products and the ejector mass flow rate distribution of the aircraft surface to the flow field wall grid points to obtain the boundary conditions for the aerodynamic thermal environment calculation.

[0013] Step Six: Determine the time based on the simulated duration in Step Four, determine the flight trajectory based on this time, and determine the flight Mach number, altitude, angle of attack, and sideslip angle based on the flight trajectory. Input the aircraft surface temperature distribution, aircraft surface ablation reaction product composition and ejector mass flow rate distribution, and aircraft surface internal pyrolysis reaction product composition and ejector mass flow rate distribution as mass ejector momentum boundary conditions, and perform aerodynamic thermal environment calculation based on the flow field grid to obtain the calculated aerodynamic thermal environment results.

[0014] Step 7: Determine whether the flight trajectory has been calculated. If it has, the calculation ends; otherwise, return to Step 2 and continue to solve the aerodynamic and thermal environment along the flight trajectory points.

[0015] In this context, the interaction between the aerodynamic thermal environment and the structural thermal response of the aircraft, and the significant impact of mass ejection behavior on the aerodynamic thermal environment, leads to changes in the aerodynamic thermal / mechanical loads on the materials. Consequently, the temperature distribution of the materials and the flow of porous media alter these changes. Furthermore, the generation, conditions, and transport processes of ablation pyrolysis gases depend on these changes. Therefore, the coupled thermal effect of mass ejection on the aerodynamic thermal / thermal response of the ablation thermal protection system cannot be ignored. For coupled simulations, it is crucial to fully consider the key processes of each physics field and the transfer of key parameters between them. Additionally, the timescale of the simulations for each physics field and the frequency of data exchange between them must also be considered.

[0016] Therefore, this invention designs a gas-thermal / thermal response coupling calculation method that includes aerodynamic thermal environment calculation under mass ejection conditions; thermal response calculation considering surface ablation reaction, internal pyrolysis reaction and pyrolysis gas flow, solid heat conduction; and a coupling data transfer method strategy between the two, considering the ejection of ablation pyrolysis products.

[0017] The coupled calculation method proposed in this invention can describe the process of gas generated by the thermal reaction of pyrolysis materials flowing through porous media to reach the surface and the gas generated by the thermal reaction of the ablation material surface; as well as the behavior of these gases being ejected into the flow field at a certain mass flow rate at the surface and interacting with the mainstream to change the flow field; this coupled calculation method can obtain the coupled aerodynamic thermal / thermal response results of the ablation pyrolysis thermal protection system of high-speed (Mach number greater than 5) aircraft considering mass ejection.

[0018] Preferably, step three uses the inverse distance weighting method to interpolate between the flow field grid and the solid field grid in space, and transfers the wall heat flux distribution and pressure distribution in the aerodynamic thermal environment calculation results to the surface grid points of the structural field to obtain the boundary conditions for thermal response calculation; step five specifically involves: interacting the flow field data and structural field data at the wall position, and transferring the aircraft structure temperature distribution, aircraft surface ablation reaction, and aircraft internal pyrolysis reaction and pyrolysis gas flow obtained in step four to the flow field wall grid points through the inverse distance weighting method to obtain the boundary conditions for aerodynamic thermal environment calculation;

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

[0020] ;

[0021] ;

[0022] ;

[0023] in, Given the physical quantity at the interpolation point, For the physical quantity at the point to be interpolated, Given the distance between the interpolation point and the point to be interpolated. and For attenuation control, Let be the weight of interpolation node i, representing the contribution of the physical information at that point to the interpolation point during that time interval; n is the number of interpolation points; and x and y are the x and y coordinates of the interpolation point, respectively. i and y i These are the x and y coordinates of the known interpolation points, respectively.

[0024] Preferably, step two specifically involves: selecting a flight trajectory point, and based on the flight Mach number, altitude, angle of attack, and sideslip angle at the corresponding moment of the selected flight trajectory point, solving the three-dimensional compressible Navier-Stokes equations using the finite volume method based on the flow field grid to obtain the aerodynamic thermal environment calculation results; step six involves solving the three-dimensional compressible Navier-Stokes equations using the finite volume method, performing aerodynamic thermal environment calculations based on the flow field grid, and obtaining the calculated aerodynamic thermal environment results.

[0025] Preferably, the mass conservation equation is:

[0026] ;

[0027] in, Porosity of the material The constant of the pyrolysis gas. For temperature, For pressure, For time, For material permeability, For the pyrolysis gas velocity, The porosity of the material before pyrolysis. The density of the material before pyrolysis. The porosity of the material after pyrolysis. The density of the material after pyrolysis. The density of the solid material. This is the partial derivative operator.

[0028] The preferred energy conservation equation is:

[0029] ;

[0030] in, For the density of pyrolysis gas, The specific heat capacity of a gas at constant pressure. For temperature, For time, Porosity of the material The density of the solid material. The specific heat capacity at constant pressure of a solid. The thermal conductivity coefficient, Let Darcy's velocity be the gas velocity. The enthalpy of the gaseous products. The enthalpy of the material, It is the heat flow from pyrolysis.

[0031] Preferably, the mass ejection momentum boundary condition is:

[0032] ;

[0033] ;

[0034] ;

[0035] ;

[0036] ;

[0037] ;

[0038] in, The temperature at the surface is also the temperature of the surface mass of the ejected gas. The set temperature (in the calculation, the surface temperature distribution obtained in step four and passed in step five). and The velocities in the two illegal outward directions at the surface are: The density at the surface, The pressure at the first layer of mesh near the surface. The density at the first layer of the grid near the surface. The velocity at the first mesh layer near the surface. Let be the mass ejector gas constant. The mass flow rate of the ejected gas is the surface quality. The velocity is at the surface.

[0039] One or more technical solutions provided by this invention have at least the following technical effects or advantages:

[0040] This invention establishes a numerical calculation method for the aerodynamic thermal environment under mass ejection conditions; a numerical calculation method for the thermal response considering surface ablation reaction, internal pyrolysis reaction, pyrolysis gas flow, and solid heat conduction; and a coupled aerodynamic / thermal response coupling calculation method considering the ejection of ablation pyrolysis products, based on a coupling data transfer strategy between the two. It can simulate the generation, transport, and ejection processes of surface ablation reaction products and internal pyrolysis reaction products in the ablation pyrolysis thermal protection system of high-speed (Mach number greater than 5) aircraft. It can calculate the coupled aerodynamic / thermal response results of the high-speed (Mach number greater than 5) aircraft ablation pyrolysis thermal protection system considering mass ejection, supporting the evaluation and design of such systems. Attached Figure Description

[0041] The accompanying drawings, which are provided to further illustrate embodiments of the invention and constitute a part of this invention, are not intended to limit the scope of the invention.

[0042] Figure 1 A flowchart illustrating the calculation method for the coupling of aerodynamic heat and thermal response considering the ejection of ablation pyrolysis products;

[0043] Figure 2 This is a schematic diagram of coupled data transmission;

[0044] Figure 3 A schematic diagram showing the geometric shape and dimensions of the object being calculated;

[0045] Figure 4 This is a schematic diagram of the flow field calculation mesh;

[0046] Figure 5 This is a schematic diagram of the structural field calculation mesh;

[0047] Figure 6 This is a schematic diagram of the structural temperature distribution at 1 second.

[0048] Figure 7 This is a schematic diagram of the structural temperature distribution at time 2s;

[0049] Figure 8This is a schematic diagram of the flow field temperature distribution and near-wall streamlines at time 1s;

[0050] Figure 9 Schematic diagrams of wall heat flow considering / not considering mass ejection at times 0s, 1s, and 2s. Detailed Implementation

[0051] To better understand the above-mentioned objectives, features, and advantages of the present invention, the present invention will be further described in detail below with reference to the accompanying drawings and specific embodiments. It should be noted that, where there is no conflict, the embodiments of the present invention and the features thereof can be combined with each other.

[0052] Many specific details are set forth in the following description in order to provide a full understanding of the invention. However, the invention may also be practiced in other ways different from those described herein, and therefore the scope of protection of the invention is not limited to the specific embodiments disclosed below.

[0053] Example 1;

[0054] Please refer to Figures 1-2 , Figure 1 A flowchart illustrating the coupled calculation method for aerodynamic heat and thermal response considering the ejection of ablation pyrolysis products. Figure 2 This is a schematic diagram of coupled data transmission. Figure 2 In this paper, p represents pressure, q represents heat flux, T represents temperature, c represents composition, subscript w represents wall surface, subscript py represents pyrolysis gas, and subscript ab represents ablation gas. This invention provides a coupled aerodynamic / thermal response calculation method considering the ejection of ablation pyrolysis products. The flow field thermal environment is calculated by solving the three-dimensional compressible Navier-Stokes equations, the structural temperature field is solved by solving the three-dimensional heat conduction equations, and the generation, transport, and reaction processes of the ejected mass gas are simulated by incorporating multi-component gas mass ejection momentum boundary conditions, a high-temperature gas chemical reaction model, Darcy's law for porous media, and an ablation pyrolysis reaction model. The solution is coupled through an established inter-field data transfer method.

[0055] The specific steps of the coupled aerodynamic / thermal response calculation method considering the ejection of ablation pyrolysis products are as follows:

[0056] Step 1: Determine the geometry of the object to be calculated and draw the flow field mesh and the structural field mesh. The object to be calculated is usually an aircraft.

[0057] Step 2: Select the flight trajectory point. Based on the flight Mach number, altitude, angle of attack, and sideslip angle at that moment (corresponding to the selected flight trajectory point), solve for the aerodynamic and thermal environment using the flow field mesh. In this step, the wall is set as a viscous solid wall boundary condition, and the wall temperature is selected as needed. Solve the three-dimensional compressible Navier-Stokes equations using the finite volume method.

[0058] The aerodynamic thermal environment calculation results under these conditions were obtained. The three-dimensional compressible Navier-Stokes equations can be found in Gui Yewei, Tang Wei, and Du Yanxia. Thermal Safety of Near-Space Hypersonic Vehicles [M]. Beijing: National Defense Industry Press, 2019: pp. 24-26. Further details are omitted in this embodiment.

[0059] Step 3: Interact the flow field data and structural field data at the wall location, transferring the wall heat flux distribution and pressure distribution from the aerodynamic thermal environment calculation results to the structural field surface grid points to establish boundary conditions for thermal response calculation. Then, use the inverse distance weighting method to interpolate between the flow field grid and the solid field grid in space. The inverse distance weighting method calculation formula is as follows:

[0060] ;

[0061] ;

[0062] ;

[0063] in, Given the physical quantity at the interpolation point, For the physical quantity at the point to be interpolated, Given the distance between the interpolation point and the point to be interpolated. and For attenuation control, it is generally taken as... , . Let be the weight of interpolation node i, representing the contribution of the physical information at that point to the point to be interpolated during that time interval. Let n be the number of interpolation points, and x and y be the x and y coordinates of the point to be interpolated, respectively. i and y i These are the x and y coordinates of the known interpolation points, respectively.

[0064] Step 4: Taking into account the surface ablation reaction, internal pyrolysis, and pyrolysis gas flow, a mass conservation equation can be established:

[0065] ;

[0066] Taking into account the heat absorption of the ablation reaction on the material surface and the internal pyrolysis reaction, the heat conduction of the solid phase, and the convective heat transfer of the pyrolysis gas, an energy conservation equation can be established:

[0067] ;

[0068] Using the heat flow distribution and pressure distribution transmitted in step three as input boundary conditions, thermal response calculations are performed based on the structural field mesh and material properties using the mass conservation equation and energy conservation equation. The simulation is conducted for several durations (which can be selected as needed) to obtain the temperature rise inside the structure, the surface ablation reaction, the internal pyrolysis reaction, and the flow of pyrolysis gas.

[0069] Step 5: Interact the flow field data and structural field data at the wall location. Transfer the surface temperature distribution, surface ablation reaction product composition and ejector mass flow rate distribution, and internal pyrolysis reaction product composition and ejector mass flow rate distribution from the thermal response calculation results in Step 4 to the flow field wall grid points through the inverse distance weighting method in Step 3 to serve as the boundary conditions for aerodynamic thermal environment calculation.

[0070] Step Six: Determine the time based on the simulation duration in Step Four, and determine the flight Mach number and altitude by combining the flight trajectory. Input the surface temperature distribution, surface ablation reaction product composition, ejector mass flow rate distribution, internal pyrolysis reaction product composition, and ejector mass flow rate from Step Four as mass ejector momentum boundary conditions. Solve the three-dimensional compressible Navier-Stokes equations using the finite volume method from Step One, and perform aerodynamic thermal environment calculations based on the flow field mesh to obtain the calculation results. The mass ejector momentum boundary conditions are as follows:

[0071] ;

[0072] ;

[0073] ;

[0074] ;

[0075] ;

[0076] ;

[0077] Substitute the ejection conditions for different gas components into the above equation.

[0078] Step 7: Determine whether the flight trajectory has been calculated. If it has, the calculation ends; otherwise, return to Step 2 and continue to solve the aerodynamic and thermal environment along the flight trajectory points.

[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] Table 2 Material properties after pyrolysis

[0094] 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 Perform inverse distance weighting calculation, interpolate the thermal response calculation results at the gas-solid interface from the structural field grid points to the flow field grid points, and generate the file Solid-Flow_1s.dat. This file contains the surface temperature distribution, surface ablation reaction product composition and ejector mass flow rate distribution, and internal pyrolysis reaction product composition and ejector mass flow rate distribution at the surface obtained in step four. The inverse distance weighting calculation formula is shown in step three.

[0095] Step Six: Based on the simulated duration of 1 second in Step Three, determine the flight time as 1 second. Combined with the flight trajectory, determine the flight Mach number as 11, flight altitude as 50 km, angle of attack as 5°, and sideslip angle as 0°. Input the surface temperature distribution, surface ablation reaction product composition, ejector mass flow rate distribution, and internal pyrolysis reaction product composition and ejector mass flow rate from Step Five as mass ejection momentum boundary conditions. Solve the three-dimensional compressible Navier-Stokes equations using the finite volume method from Step Two, based on... Figure 4 The aerodynamic thermal environment was calculated using the flow field mesh, and the results were obtained. The mass ejection momentum boundary condition can be expressed as:

[0096] ;

[0097] ;

[0098] ;

[0099] ;

[0100] ;

[0101] ;

[0102] Substitute the conditions for different ejector gas components into the above formula to determine the overall ejection conditions.

[0103] Step 7: Perform the same operation as in Step 3 on the aerodynamic and thermal environment results obtained in Step 6. Repeat Steps 3 and 6, calculating along the flight trajectory until completion. In this example, the calculation ends at 2 seconds, yielding the aerodynamic and thermal response results at the required time along the flight trajectory. At 2 seconds, the corresponding flight trajectory yields the following flight conditions: Mach number 11, altitude 48 km, angle of attack 8°, and sideslip angle 0°.

[0104] Figures 6-7 The structural temperature distributions at times 1s and 2s are shown; Figure 8 The flow field temperature distribution and near-wall streamlines at time 1 second are shown; Figure 9 The wall heat flux results at times 0s, 1s, and 2s are presented. Figure 9The horizontal axis represents the distance in the X direction, in millimeters, and the vertical axis represents the wall heat flux, in kilowatts per square meter. Legend: 0s is the value at time 0s of the coupling calculation; 1s-NoInject is the value at time 1s of the coupling calculation without considering ejection; 1s-Inject is the value at time 1s of the coupling calculation with considering ejection; 2s-NoInject is the value at time 2s of the coupling calculation without considering ejection; 2s-Inject is the value at time 2s of the coupling calculation with considering ejection.

[0105] Although preferred embodiments of the invention have been described, those skilled in the art, upon learning the basic inventive concept, can make other changes and modifications to these embodiments. Therefore, the appended claims are intended to be interpreted as including both the preferred embodiments and all changes and modifications falling within the scope of the invention.

[0106] Obviously, those skilled in the art can make various modifications and variations to this invention without departing from its spirit and scope. Therefore, if these modifications and variations fall within the scope of the claims of this invention and their equivalents, this invention also intends to include these modifications and variations.

Claims

1. A method for coupled calculation of aerodynamic heat and thermal response considering the ejection of ablation pyrolysis products, characterized in that, The method includes: Step 1: Determine the geometric shape of the aircraft and draw the flow field mesh and structural field mesh; Step 2: Select the flight trajectory point. Based on the flight Mach number, altitude, angle of attack and sideslip angle at the corresponding moment of the selected flight trajectory point, solve the aerodynamic thermal environment based on the flow field grid to obtain the aerodynamic thermal environment calculation results. Step 3: Interact the flow field data and structural field data at the wall location, and transfer the wall heat flux distribution and pressure distribution in the aerodynamic thermal environment calculation results to the grid points on the structural field surface to obtain the boundary conditions for thermal response calculation; Step 4: Establish mass conservation equations considering the ablation reaction on the aircraft surface, internal pyrolysis, and pyrolysis gas flow; establish energy conservation equations considering the heat absorption of the ablation reaction on the aircraft material surface and the internal pyrolysis reaction, the heat conduction of the solid phase, and the convective heat transfer of the pyrolysis gas. Based on the boundary conditions for thermal response calculation obtained in Step 3, perform thermal response calculations using the mass conservation equations and energy conservation equations, simulate for a set duration, and obtain the temperature distribution of the aircraft structure, the ablation reaction on the aircraft surface, and the internal pyrolysis reaction and pyrolysis gas flow. Step 5: Interact the flow field data and structural field data at the wall location. Obtain the surface temperature distribution of the aircraft based on the structural temperature distribution of the aircraft. Obtain the composition of the ablation reaction products and the ejector mass flow rate distribution of the aircraft surface based on the ablation reaction of the aircraft surface. Obtain the composition of the internal pyrolysis reaction products and the ejector mass flow rate distribution of the aircraft surface based on the internal pyrolysis reaction and pyrolysis gas flow of the aircraft. Transfer the surface temperature distribution, the composition of the ablation reaction products and the ejector mass flow rate distribution of the aircraft surface, and the composition of the internal pyrolysis reaction products and the ejector mass flow rate distribution of the aircraft surface to the flow field wall grid points to obtain the boundary conditions for the aerodynamic thermal environment calculation. Step Six: Determine the time based on the simulated duration in Step Four, determine the flight trajectory based on this time, and determine the flight Mach number, altitude, angle of attack, and sideslip angle based on the flight trajectory. Input the aircraft surface temperature distribution, aircraft surface ablation reaction product composition and ejector mass flow rate distribution, and aircraft surface internal pyrolysis reaction product composition and ejector mass flow rate distribution as mass ejector momentum boundary conditions, and perform aerodynamic thermal environment calculation based on the flow field grid to obtain the calculated aerodynamic thermal environment results. Step 7: Determine whether the flight trajectory has been calculated. If it has, the calculation ends; otherwise, return to Step 2 and continue to solve the aerodynamic and thermal environment along the flight trajectory points.

2. The aerodynamic and thermal response coupling calculation method considering the ejection of ablation pyrolysis products according to claim 1, characterized in that, Step three involves interpolating between the flow field grid and the solid field grid in space using the inverse distance weighting method. This interpolates the wall heat flux distribution and pressure distribution from the aerodynamic thermal environment calculation results to the surface grid points of the structural field, thereby obtaining the boundary conditions for thermal response calculation. Step five specifically involves interacting the flow field data and structural field data at the wall location. 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 obtained in step four are then transmitted to the flow field wall grid points using the inverse distance weighting method, thereby obtaining the boundary conditions for aerodynamic thermal environment calculation. The inverse distance weighting method calculation formula is: ； ； ； in, Given the physical quantity at the interpolation point, For the physical quantity at the point to be interpolated, Given the distance between the interpolation point and the point to be interpolated. and For attenuation control, Let be the weight of interpolation node i, n be the number of interpolation points, and x and y be the x and y coordinates of the point to be interpolated, respectively. i and y i These are the x and y coordinates of the known interpolation points, respectively.

3. The aerodynamic and thermal response coupling calculation method considering the ejection of ablation pyrolysis products according to claim 1, characterized in that, Step two specifically involves: selecting a flight trajectory point, and based on the flight Mach number, altitude, angle of attack, and sideslip angle at the corresponding moment of the selected flight trajectory point, solving the three-dimensional compressible Navier-Stokes equations using the finite volume method based on the flow field grid to obtain the aerodynamic thermal environment calculation results; Step six involves solving the three-dimensional compressible Navier-Stokes equations using the finite volume method, performing aerodynamic thermal environment calculations based on the flow field grid, and obtaining the calculated aerodynamic thermal environment results.

4. The aerodynamic and thermal response coupling calculation method considering the ejection of ablation pyrolysis products according to claim 1, characterized in that, The mass conservation equation is: ； in, Porosity of the material The constant of the pyrolysis gas. For temperature, For pressure, For time, For material permeability, For the pyrolysis gas velocity, The porosity of the material before pyrolysis. The density of the material before pyrolysis. The porosity of the material after pyrolysis. The density of the material after pyrolysis. The density of the solid material. This is the partial derivative operator.

5. The aerodynamic and thermal response coupling calculation method considering the ejection of ablation pyrolysis products according to claim 1, characterized in that, The energy conservation equation is: ； in, For the density of pyrolysis gas, The specific heat capacity of a gas at constant pressure. For temperature, For time, Porosity of the material The density of the solid material. The specific heat capacity at constant pressure of a solid. The thermal conductivity coefficient, Let Darcy's velocity be the gas velocity. The enthalpy of the gaseous products. The enthalpy of the material, It is the heat flow from pyrolysis.

6. The aerodynamic and thermal response coupling calculation method considering the ejection of ablation pyrolysis products according to claim 1, characterized in that, The mass ejection momentum boundary condition is: ； ； ； ； ； ； in, The temperature at the surface. For the set temperature, and The velocities in the two illegal outward directions at the surface. The density at the surface, The pressure at the first layer of the mesh near the surface. The density at the first layer of the grid near the surface. The velocity at the first mesh layer near the surface. Let be the mass ejector gas constant. The mass flow rate of the ejected gas is the surface quality. The velocity is at the surface.

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