A method, device, equipment and medium for aerodynamic thermal analysis under multiple conditions

Abstract

The present application discloses an aerodynamic thermal analysis method, device, equipment and medium under multiple conditions, which relates to the field of aerodynamic technology, including: obtaining a target high-temperature gas flow field of an aircraft, extracting wall parameter information and wall sub-layer parameter information; calculating the mass flux of each component flowing in and out of a control body, determining the conditional parameters of each wall effect condition, and obtaining a component mass conservation equation; constructing a total wall mass conservation equation based on the mass flux of each component, the component diffusion conservation condition and the catalytic conservation condition; constructing a wall energy conservation equation based on the total mass ejection flux determined by the total wall mass conservation equation, the wall parameter information and the wall sub-layer parameter information; performing mathematical operations on the component mass conservation equation, the wall total mass conservation equation and the wall energy conservation equation to obtain a wall heat flux calculation equation; substituting the current wall condition into the wall heat flux calculation equation to obtain an aerodynamic thermal analysis result, and realizing accurate aerodynamic thermal analysis under different working conditions.

Description

Technical Field

[0001] The invention relates to the technical field of aerodynamics, and in particular to an aerodynamic thermal analysis method, device, equipment and medium under multiple conditions. Background Art

[0002] When a high-speed aircraft re-enters or flies in near space, the gas in its flow field is strongly compressed and produces intense friction with the wall, resulting in severe aerodynamic heating. The temperature of the aircraft wall increases significantly, which has a significant impact on flight safety. Therefore, thermal protection is one of the key issues in high-speed flight.

[0003] Mass ejection has always been one of the main heat protection methods for aircraft. Traditional ablative heat protection materials absorb and disperse heat through wall ablation reactions and gas ejection from material pyrolysis, which is also called passive ejection technology. In recent years, with the development of heat protection technology, active mass ejection technologies such as sweat cooling and film cooling have gradually moved towards engineering applications. This method achieves the effect of cooling by actively injecting cooling fluid into the flow field. In recent years, active and passive ejection technologies have been favored by engineering units and researchers, but in current research, the issue of how to calculate wall heat flux under ejection conditions has not been clarified.

[0004] In summary, how to comprehensively consider the influence of multiple factors under the flow field conditions of a high-speed aircraft and achieve accurate aerodynamic thermal analysis is a technical problem to be solved in this field. Summary of the invention

[0005] In view of this, the purpose of the present invention is to provide a multi-condition aerodynamic thermal analysis method, device, equipment and medium, which can comprehensively consider the influence of factors under multiple conditions of the flow field in which a high-speed aircraft is located, and realize accurate aerodynamic thermal analysis. The specific scheme is as follows:

[0006] In a first aspect, the present application discloses a multi-condition aerodynamic thermal analysis method, comprising:

[0007] Acquiring a target high-temperature gas flow field of the aircraft to extract wall parameter information and wall sub-layer parameter information for evaluating wall heat flux from convergent flow field information of the target high-temperature gas flow field;

[0008] Calculating the mass flux of each component flowing into and out of the control volume according to the wall parameter information and the wall sub-layer parameter information, and determining the conditional parameters of each wall effect condition to construct a component mass conservation equation;

[0009] Constructing a total wall mass conservation equation based on the mass flux of each component, the component diffusion conservation condition and the catalytic conservation condition in the component mass conservation equation;

[0010] Constructing a wall energy conservation equation based on the total mass ejection flux determined by the wall total mass conservation equation, the wall parameter information, and the wall sub-layer parameter information;

[0011] By performing mathematical operations on the component mass conservation equation, the wall total mass conservation equation, and the wall energy conservation equation, a wall heat flux calculation equation is obtained that characterizes the heat flux composition felt by the solid structure of the aircraft under the action of each wall effect;

[0012] The current wall condition of the aircraft is substituted into the wall heat flux calculation equation to obtain the aerodynamic thermal analysis result of the aircraft under the corresponding wall condition.

[0013] Optionally, the calculating the mass flux of each component flowing into and out of the control body according to the wall parameter information and the wall sub-layer parameter information, and determining the condition parameters of each wall effect condition, includes:

[0014] Calculating the component diffusion flux of each component flowing into and out of the control body, the active ejection gas flux, the pyrolysis gas ejection flux, the catalytic reaction generation rate and the ablation reaction generation rate according to the wall parameter information and the wall sub-layer parameter information;

[0015] Determine the mass proportion of each component in the actively ejected gas and the mass proportion of each component in the pyrolysis gas under each wall effect condition.

[0016] Optionally, the wall total mass conservation equation is constructed based on the mass flux of each component in the component mass conservation equation, the component diffusion conservation condition and the catalytic conservation condition, including:

[0017] The component diffusion flux, the active ejection gas flux, the pyrolysis gas ejection flux, the catalytic reaction generation rate and the ablation reaction generation rate of each component in the component mass conservation equation are added together to obtain an initial wall total mass conservation equation;

[0018] Set the component diffusion conservation condition that the mass flux of each component diffusing into the control volume and the mass flux diffusing out of the control volume meet the conservation condition;

[0019] A catalytic conservation condition is set in which the total mass of each reactant component and the total mass of the product component satisfy the conservation condition during the wall catalytic reaction;

[0020] The total wall mass conservation equation is constructed based on the initial wall total mass conservation equation, the component diffusion conservation condition, and the catalytic conservation condition.

[0021] Optionally, the wall energy conservation equation is constructed based on the total mass ejection flux determined by the wall total mass conservation equation, the wall parameter information, and the wall sub-layer parameter information, including:

[0022] Based on the wall parameter information and the wall sub-layer parameter information, enthalpy parameters including wall component enthalpy, total wall ejected gas enthalpy, active ejected gas enthalpy, pyrolysis ejected gas enthalpy, active ejected gas generation enthalpy, pyrolysis ejected gas generation enthalpy and conduction heat flux items are calculated and determined;

[0023] A wall energy conservation equation characterizing the energy transferred from the control body to the flow field and the solid structure of the aircraft is constructed according to the conduction heat flux term, the enthalpy parameter and the total mass ejection flux determined based on the wall total mass conservation equation.

[0024] Optionally, the mathematical operation of the component mass conservation equation, the wall total mass conservation equation, and the wall energy conservation equation is performed to obtain a wall heat flux calculation equation that characterizes the heat flux composition felt by the solid structure of the aircraft under the action of each wall effect, including:

[0025] Multiplying the component mass conservation equation with the wall component enthalpy value in the wall energy conservation equation and summing the values ​​for each component to obtain a total energy change equation caused by changes in each wall mass flux;

[0026] According to the mathematical operation of the total energy change equation and the wall total mass conservation equation, a wall heat flux calculation equation is obtained to characterize the heat flux composition felt by the solid structure of the aircraft under the action of each wall effect.

[0027] Optionally, the mathematical operation based on the total energy change equation and the wall total mass conservation equation to obtain a wall heat flux calculation equation characterizing the heat flux composition felt by the solid structure of the aircraft under the action of each wall effect includes:

[0028] According to mathematical operations of the total energy change equation and the wall total mass conservation equation, a wall heat flow calculation equation constructed based on the conduction heat flux term, the wall component enthalpy value, the active ejection gas generation enthalpy and the pyrolysis ejection gas generation enthalpy is obtained.

[0029] Optionally, obtaining a target high-temperature gas flow field of an aircraft includes:

[0030] The target high-temperature gas flow field of the aircraft is calculated using a preset high-temperature gas thermochemical non-equilibrium flow field numerical simulation method.

[0031] In a second aspect, the present application discloses a multi-condition pneumatic thermal analysis device, comprising:

[0032] An information acquisition module, used to acquire a target high-temperature gas flow field of the aircraft, so as to extract wall parameter information and wall sub-layer parameter information for evaluating wall heat flux from convergent flow field information of the target high-temperature gas flow field;

[0033] A first equation building module is used to calculate the mass flux of each component flowing into and out of the control volume according to the wall parameter information and the wall sub-layer parameter information, and determine the condition parameters of each wall effect condition to construct a component mass conservation equation;

[0034] A second equation building module is used to build a wall total mass conservation equation based on the mass flux of each component, the component diffusion conservation condition and the catalytic conservation condition in the component mass conservation equation;

[0035] A third equation building module is used to build a wall energy conservation equation based on the total mass ejection flux determined by the wall total mass conservation equation, the wall parameter information, and the wall sub-layer parameter information;

[0036] A heat flow equation determination module is used to obtain a wall heat flow calculation equation that characterizes the heat flow composition felt by the solid structure of the aircraft under the action of each wall effect by performing mathematical operations on the component mass conservation equation, the wall total mass conservation equation, and the wall energy conservation equation;

[0037] The analysis module is used to substitute the current wall condition of the aircraft into the wall heat flux calculation equation to obtain the aerodynamic thermal analysis result of the aircraft under the corresponding wall condition.

[0038] In a third aspect, the present application discloses an electronic device, comprising:

[0039] Memory, used to store computer programs;

[0040] A processor is used to execute the computer program to implement the steps of the aerodynamic thermal analysis method under multiple conditions disclosed above.

[0041] In a fourth aspect, the present application discloses a computer-readable storage medium for storing a computer program; wherein, when the computer program is executed by a processor, the steps of the aerodynamic thermal analysis method under multiple conditions disclosed above are implemented.

[0042] It can be seen that the present application discloses an aerodynamic thermal analysis method under multiple conditions, including: obtaining a target high-temperature gas flow field of an aircraft to extract wall parameter information and wall sublayer parameter information for evaluating wall heat flux from the convergent flow field information of the target high-temperature gas flow field; calculating the mass flux of each component flowing into and out of the control body according to the wall parameter information and the wall sublayer parameter information, and determining the conditional parameters of each wall effect condition to construct a component mass conservation equation; constructing a component mass conservation equation based on the mass flux of each component, the component diffusion conservation condition and the catalytic conservation condition in the component mass conservation equation. The total mass conservation equation of the wall; the wall energy conservation equation is constructed based on the total mass ejection flux determined by the total mass conservation equation of the wall, the wall parameter information, and the wall sub-layer parameter information; through mathematical operations on the component mass conservation equation, the total mass conservation equation of the wall, and the wall energy conservation equation, a wall heat flow calculation equation is obtained to characterize the heat flow composition of the solid structure of the aircraft under the action of each wall effect; the current wall condition of the aircraft is substituted into the wall heat flow calculation equation to obtain the aerodynamic thermal analysis results of the aircraft under the corresponding wall condition. It can be seen that by constructing the component mass conservation equation, the total mass conservation equation of the wall, and the wall energy conservation equation in detail, and obtaining the wall heat flow calculation equation based on this, the wall heat flow can be accurately calculated. This scheme comprehensively considers a variety of wall effect conditions, including wall catalysis, ablation, pyrolysis ejection, active ejection, etc., so that the wall heat flow calculation equation is applicable to wall conditions of different complexity. Whether it is a simple conventional catalytic wall without ablation or ejection, or a complex wall with multiple effects, this method can be used to perform aerodynamic thermal analysis, providing an accurate evaluation method for the thermal protection research of aircraft under various working conditions. BRIEF DESCRIPTION OF THE DRAWINGS

[0043] In order to more clearly illustrate the embodiments of the present invention or the technical solutions in the prior art, the drawings required for use in the embodiments or the description of the prior art will be briefly introduced below. Obviously, the drawings described below are only embodiments of the present invention. For ordinary technicians in this field, other drawings can be obtained based on the provided drawings without paying creative work.

[0044] Figure 1 A flow chart of a multi-condition aerodynamic thermal analysis method disclosed in this application;

[0045] Figure 2 This is a schematic diagram of a component mass conservation equation constructed based on various components and multiple wall effect conditions disclosed in this application;

[0046] Figure 3 A schematic diagram of constructing a wall energy conservation equation disclosed in this application;

[0047] Figure 4 A simplified model schematic diagram of an aircraft disclosed in the present application;

[0048] Figure 5 A wall heat flux distribution diagram disclosed in this application;

[0049] Figure 6 Another wall heat flux distribution diagram disclosed in this application;

[0050] Figure 7 This is a schematic diagram of the structure of a pneumatic thermal analysis device under multiple conditions disclosed in this application;

[0051] Figure 8 This is a structural diagram of an electronic device disclosed in this application. DETAILED DESCRIPTION

[0052] The following will be combined with the drawings in the embodiments of the present application to clearly and completely describe the technical solutions in the embodiments of the present application. Obviously, the described embodiments are only part of the embodiments of the present invention, not all of the embodiments. Based on the embodiments of the present invention, all other embodiments obtained by ordinary technicians in this field without making creative work are within the scope of protection of the present invention.

[0053] When a high-speed aircraft re-enters or flies in near space, the gas in its flow field is strongly compressed and produces intense friction with the wall, resulting in severe aerodynamic heating. The temperature of the aircraft wall increases significantly, which has a significant impact on flight safety. Therefore, thermal protection is one of the key issues in high-speed flight.

[0054] Mass ejection has always been one of the main heat protection methods for aircraft. Traditional ablative heat protection materials absorb and disperse heat through wall ablation reactions and gas ejection from material pyrolysis, which is also called passive ejection technology. In recent years, with the development of heat protection technology, active mass ejection technologies such as sweat cooling and film cooling have gradually moved towards engineering applications. This method achieves the effect of cooling by actively injecting cooling fluid into the flow field. In recent years, active and passive ejection technologies have been favored by engineering units and researchers, but in current research, the issue of how to calculate wall heat flux under ejection conditions has not been clarified.

[0055] For the numerical simulation of the complete gas flow field, the wall heat flux Only conduction heat flow terms are included, namely:

[0056] ;

[0057] in, is the thermal conductivity of gas, represents the wall surface, is the wall temperature Along the wall normal direction The gradient of the wall temperature.

[0058] However, for the thermochemical non-equilibrium flow of high-speed aircraft, due to the severe aerodynamic heating, the temperature in the shock layer flow field is very high, which will lead to thermochemical reactions such as vibration and bound electron excitation, dissociation and ionization, and produce transport phenomena such as convection and component diffusion as the high-speed flow field moves. At the same time, the catalytic effect of the wall material will also cause the gas components to undergo a complex reaction on the wall. Therefore, the wall heat flux no longer only includes the above-mentioned conduction heat flux term, but also includes the diffusion (catalytic) heat flux term brought by the component transport. Taking the numerical simulation of the chemical non-equilibrium-temperature model as an example, the wall heat flux calculation formula is as follows:

[0059] ;

[0060] The first term on the right side of the above formula is the conduction heat flow term, which is consistent with the perfect gas model. is the component diffusion heat, where: is the gas density, , and The components are The mass fraction, diffusion coefficient and enthalpy of represents the wall, is the component number, is the number of components, is the mass fraction gradient of the wall component.

[0061] For walls with ablation and pyrolysis ejection, the source of the wall heat flux becomes complicated. If the above formula is still used, the evaluation of the overall aerodynamic heat is not comprehensive enough. For high-temperature thermochemical nonequilibrium flows with active and passive ejection, if the above formula is still used to evaluate the wall heat flux, it will bring results that are contrary to the actual situation. For example, if water is introduced as the main working fluid, at a certain section of the aircraft wall, due to the large amount of water ejected, it immediately causes a huge mass fraction reverse gradient near the wall, thereby introducing a huge diffusion heat flux. At this time, if the above formula is still used to evaluate the wall heat flux, it will be concluded that mass ejection actually increases the wall heat flux, but this is obviously inconsistent with the actual mass ejection heat reduction situation. Therefore, under the conditions of active and passive ejection, the wall heat flux calculation can no longer use the above formula, but a new evaluation method must be adopted.

[0062] To this end, the present invention provides an aerodynamic thermal analysis scheme under multiple conditions, which can comprehensively consider the influence of factors under multiple conditions in the flow field where a high-speed aircraft is located, and realize accurate aerodynamic thermal analysis.

[0063] Reference Figure 1As shown, the embodiment of the present invention discloses a method for aerodynamic thermal analysis under multiple conditions, including:

[0064] Step S11: Acquire a target high-temperature gas flow field of the aircraft to extract wall parameter information and wall sub-layer parameter information for evaluating wall heat flux from convergent flow field information of the target high-temperature gas flow field.

[0065] In this embodiment, a preset high-temperature gas thermochemical non-equilibrium flow field numerical simulation method is used to calculate the target high-temperature gas flow field of the aircraft. It can be understood that for any high-speed aircraft, during high-speed flight, the friction between the aircraft and the air will produce a significant thermal effect, causing the temperature of the surrounding gas flow field to rise sharply. This high-temperature environment may trigger reactions such as catalysis and ablation, which have an important impact on the material and structural properties of the aircraft. The usual high temperature range is 1000K to 30000K. The gas flow field within this range is the high-temperature gas flow field. The conventional high-temperature gas thermochemical non-equilibrium flow field numerical simulation method is used to calculate the target high-temperature gas flow field. Among them, the steps for calculating the target high-temperature gas flow field are as follows:

[0066] Establishment of basic equations: First, it is necessary to establish basic equations applicable to high-temperature gas flow fields, including the mass conservation equation, momentum conservation equation, and energy conservation equation. Under thermochemical non-equilibrium conditions, these equations will take into account factors such as changes in gas components and energy release and absorption caused by chemical reactions.

[0067] Chemical reaction model selection: Select a chemical reaction model to describe the chemical reaction process in high-temperature gas. Since the gas temperature is high in hypersonic flow, various chemical reactions such as dissociation and recombination will occur, and an accurate chemical reaction model is required to reflect these processes. For example, a multi-component chemical reaction model that considers multiple components and reactions may be used.

[0068] Numerical discretization method: numerically discretize the established basic equations and chemical reaction models. This usually involves converting continuous partial differential equations into discrete algebraic equations so that they can be solved by computer. Common numerical discretization methods include finite difference method and finite volume method, which divide the computational domain into multiple grid cells and perform approximate calculations on the physical quantities in each cell.

[0069] Boundary condition setting: Set appropriate boundary conditions. For high-temperature gas flow field calculations, boundary conditions include incoming flow conditions (such as incoming flow velocity, temperature, composition, etc.) and wall conditions (such as wall temperature, wall catalytic properties, ejection conditions, etc.). The accurate setting of these boundary conditions is crucial to obtaining the correct flow field solution.

[0070] Iterative solution: The discretized equations are solved using an iterative algorithm. Starting from the initial guess of the flow field solution, the physical quantities in the flow field (such as velocity, temperature, component concentration, etc.) are continuously updated iteratively until the convergence conditions are met to obtain the target high-temperature gas flow field. During the iteration process, the gas components and the corresponding chemical reaction source terms are continuously updated according to the chemical reaction model.

[0071] In this embodiment, based on the convergent flow field information of the target high-temperature gas flow field, the wall parameter information and the wall sub-layer parameter information are extracted, so that the extracted wall parameter information can be used for equation construction of the subsequent aerodynamic thermal analysis process.

[0072] Step S12: Calculate the mass flux of each component flowing into and out of the control volume according to the wall parameter information and the wall sub-layer parameter information, and determine the conditional parameters of each wall effect condition to construct a component mass conservation equation.

[0073] In this embodiment, the component diffusion flux, active ejection gas flux, pyrolysis gas ejection flux, catalytic reaction generation rate and ablation reaction generation rate of each component flowing into and out of the control body are calculated based on the wall parameter information and the wall sub-layer parameter information; the mass proportion of each component in the active ejection gas and the mass proportion of each component in the pyrolysis gas under each wall effect condition are determined. It can be understood that for a control body including a gas flow field, solid structure and interface near the wall of an aircraft, after considering the ablation reaction, gas pyrolysis, active mass ejection and catalytic effect of the wall, the mass flux of the component flowing into and out of the control body is distinguished, and the component mass conservation condition can be expressed as: the total mass flux of component i entering the flow field from the wall, including the component diffusion flux, active ejection gas flux, pyrolysis gas ejection flux, catalytic reaction generation rate and ablation reaction generation rate. The obtained component mass conservation equation is as follows:

[0074] ;Formula (1)

[0075] in, Indicates the component number, , is the total number of components; represents the wall surface, is the total wall mass ejection flux, , and For components Mass fraction, mass fraction gradient and diffusion coefficient at the wall, represents the normal vector perpendicular to the wall, is the wall gas density; Indicates active ejection. is the active ejection gas mass flux, is the mass proportion of component i in the actively ejected gas; represents pyrolysis induced, is the pyrolysis gas ejection mass flux, The components in the pyrolysis gas The mass percentage of represents the wall catalytic reaction, Components caused by wall catalytic reaction The rate of change of mass per unit area, i.e., the rate of generation of the catalytic reaction; represents the wall ablation reaction, is the mass change rate per unit area of ​​component i caused by the ablation reaction, that is, the ablation reaction generation rate. , , , , It is calculated based on the wall parameter information and the wall sub-layer parameter information. In addition, the current It cannot be determined yet and needs to be determined according to the following steps.

[0076] Figure 2 is a schematic diagram of a physical model related to wall heat flux and component mass flux. Figure 2 It mainly includes the following parts: wall (Solid) and gas (Gas) Figure 2 There is a horizontal dotted line in the middle representing the wall. The gas area is above the wall, and the solid area is below the wall.

[0077] A= , this expression represents the component diffusion flux of the component at the wall, B = , which represents the product of the total mass ejection flux at the wall and the mass fraction of the component at the wall.

[0078] C= and D = : represent the product of the mass flux of active gas injection and pyrolysis gas injection and the mass fraction of the corresponding component.

[0079] E= and F = : represent the rate of change of component mass per unit area caused by wall catalytic reaction and wall ablation reaction respectively.

[0080] Figure 2The arrows in indicate the direction of each physical quantity, where an arrow pointing outward indicates outflow from the control volume, and an arrow pointing inward indicates inflow into the control volume. For example, an arrow pointing from the wall to the gas indicates that the mass flux flows from the wall to the gas, while an arrow pointing from the gas to the wall indicates that physical quantities such as diffusion flux flow from the gas to the wall. Through intuitive symbols and arrows, the directions and relationships of the mass fluxes of components involved in different physical processes (such as diffusion, ejection, catalytic reaction, ablation reaction, etc.) at the interface between the wall and the gas are shown.

[0081] Step S13: constructing a total wall mass conservation equation based on the mass flux of each component, the component diffusion conservation condition and the catalytic conservation condition in the component mass conservation equation.

[0082] In this embodiment, the component diffusion flux, the active ejection gas flux, the pyrolysis gas ejection flux, the catalytic reaction generation rate and the ablation reaction generation rate of each component in the component mass conservation equation are added to obtain the initial wall total mass conservation equation; the component diffusion conservation condition is set so that the mass flux of each component diffused into the control body and the mass flux diffused out of the control body meet the conservation condition; the catalytic conservation condition is set so that the total mass of each reactant component and the total mass of the product component meet the conservation condition during the wall catalytic reaction; the wall total mass conservation equation is constructed based on the initial wall total mass conservation equation, the component diffusion conservation condition and the catalytic conservation condition. It can be understood that the Ns components in the above formula (1) are added to obtain the initial wall total mass conservation equation, which is as follows:

[0083] ;Formula (2)

[0084] At the same time, the component diffusion conservation condition and catalytic conservation condition are introduced, namely:

[0085] ;Formula (3)

[0086] ;Formula (4)

[0087] According to formula (2), formula (3) and formula (4), the total wall mass conservation equation is obtained as follows:

[0088] ;Formula (5)

[0089] In this way, the total wall mass conservation equation can be understood as the total mass flux entering the flow field from the wall, including the mass flux of actively induced gas, the mass flux of pyrolysis gas, and the mass flux of solid wall material entering the flow field caused by wall ablation reaction. Because the diffusion and catalytic effects are conserved on the wall, they will not cause changes in the total wall mass flux.

[0090] Step S14: constructing a wall energy conservation equation based on the total mass ejection flux determined by the wall total mass conservation equation, the wall parameter information, and the wall sub-layer parameter information.

[0091] In this embodiment, based on the wall parameter information and the wall sub-layer parameter information, enthalpy parameters including wall component enthalpy, total wall induced gas enthalpy, active induced gas enthalpy, pyrolysis induced gas enthalpy, active induced gas generation enthalpy, pyrolysis induced gas generation enthalpy and conduction heat flux terms are calculated and determined; according to the conduction heat flux terms, the enthalpy parameters and the total mass induced flux determined based on the wall total mass conservation equation, a wall energy conservation equation is constructed to characterize the energy transferred from the control body to the flow field and the solid structure of the aircraft. It can be understood that for the control body including the gas flow field, solid structure and interface near the wall of the aircraft, the various wall effects mentioned above are taken into account, and the heat flux conducted to the solid structure of the aircraft by the wall is set. , the heat flux can be dynamically adjusted to maintain the wall at the current isothermal wall condition. Differentiating the energy flux flowing into and out of the control body, the energy conservation condition can be expressed as: the energy transferred by the control body to the flow field and solid structure is balanced with the flow field conduction heat flux, diffusion energy flux, active ejection gas energy flux, and pyrolysis gas energy flux. The wall energy conservation equation is as follows:

[0092] ;Formula (6)

[0093] in, , and They represent the total wall-injected gas enthalpy, the actively-injected gas enthalpy and the pyrolysis-injected gas enthalpy, respectively. Indicates the average enthalpy change in the active injection pre-process (active injection gas generation enthalpy), such as latent heat of vaporization; Indicates the average enthalpy change in the pre-pyrolysis process (pyrolysis induced gas generation enthalpy), such as pyrolysis latent heat; Indicates components Enthalpy of the wall component; is the conduction heat flux term of the flow field to the wall.

[0094] in, and It is usually given in some form. The wall state parameters in the wall parameter information and the wall sub-layer parameter information determined by the calculation in the aforementioned step S11 are calculated.

[0095] , and According to the wall component enthalpy The sum of the content of each component can be obtained, and the calculation method is as follows:

[0096] ;Formula (7)

[0097] Conduction heat flux at the wall The calculation method is:

[0098] ;Formula (8)

[0099] in, and are the thermal conductivity and temperature gradient of the flow field gas on the wall, respectively.

[0100] It should be noted that in formula (6) The parameters are calculated according to formula (2), and all other parameters are calculated from the parameter information extracted in step S11.

[0101] like Figure 3 Schematic diagram of the physical model involving heat flow and mass flux at the gas-solid interface. Figure 3 There is a horizontal dotted line in the figure that represents the interface between the solid and the gas. Above the interface is the gas region, and below is the solid region. In the gas region above the interface, there are multiple symbols representing different physical quantities.

[0102] G= , which represents the sum of the products of the diffusion flux and enthalpy of each component at the wall, H = represents the product of the total mass flux at the wall and the wall enthalpy, I= represents a heat flux from the gas to (conducted to) the wall.

[0103] The physical quantities in the solid region are expressed as follows:

[0104] J= K = represents the product of the mass flux of the actively ejected gas, the sum of the ejected gas enthalpy and the ejected gas generation enthalpy at the wall. It represents the product of the mass flux of pyrolysis gas ejection and the sum of the enthalpy of pyrolysis gas and the enthalpy of pyrolysis gas formation, L= represents a heat flux directed from the wall to (conducted into) the solid.

[0105] Figure 3 The arrows in the figure indicate the direction of each physical quantity. Specifically, the arrow pointing from the wall to the gas indicates that the mass flux or heat flux flows from the wall to the gas, while the arrow pointing from the gas to the wall indicates that the diffusion flux and other physical quantities flow from the gas to the wall. In general, Figure 3The directions and relationships of the component mass fluxes, enthalpy, and heat fluxes involved in different physical processes (such as diffusion, ejection, heat flux, etc.) at the wall and gas interface are shown.

[0106] Step S15: By performing mathematical operations on the component mass conservation equation, the wall total mass conservation equation, and the wall energy conservation equation, a wall heat flux calculation equation is obtained that characterizes the heat flux composition felt by the solid structure of the aircraft under the action of each wall effect.

[0107] In this embodiment, the component mass conservation equation is multiplied with the wall component enthalpy value in the wall energy conservation equation and the sum is taken for each component to obtain the total energy change equation caused by the change of each mass flux on the wall; based on the mathematical operation of the total energy change equation and the wall total mass conservation equation, a wall heat flux calculation equation is obtained that characterizes the heat flux composition felt by the solid structure of the aircraft under the action of each wall effect. Specifically, based on the mathematical operation of the total energy change equation and the wall total mass conservation equation, a wall heat flux calculation equation is obtained that is constructed based on the conduction heat flux term, the wall component enthalpy value active induced gas generation enthalpy and the pyrolysis induced gas generation enthalpy. It can be understood that by combining formula (1) with the wall component enthalpy value Multiplying and summing all components yields formula (9), as follows:

[0108] ;Formula (9)

[0109] Finally, combining formula (6) and formula (9), we can get the wall heat flux calculation equation of the heat flux felt by the solid structure of the actual aircraft:

[0110] ;Formula (10)

[0111] In addition, the wall catalysis and ablation reaction processes are unified into the wall reaction, using The mass flux is expressed uniformly, and the wall heat flux calculation equation after arrangement is as follows:

[0112] ;Formula (11)

[0113] In this way, formula (11) is applicable to all processes involving active and passive ejection, wall ablation and catalytic reaction, and can be expressed as follows: the heat flux felt by the solid structure of the aircraft is composed of three terms: the conduction heat flux term, the wall reaction enthalpy change and the enthalpy change of the active and passive ejection pre-process.

[0114] It should be noted that formula (11) can be flexibly simplified according to the actual aerodynamic thermal analysis situation. Specifically, when the previous physical and chemical changes are not considered, that is, and , the simplified result of formula (11) is as follows:

[0115] ;Formula (12)

[0116] Formula (12) is expressed as follows: The wall heat flux felt by the solid structure of the aircraft includes the heat conduction term of the flow field and the enthalpy change caused by the wall chemical reaction effect.

[0117] For non-catalytic walls with ablation reactions, the simplified result of formula (11) is as follows:

[0118] ;Formula (13)

[0119] For a conventional catalytic wall without ablation and ejection, the simplified result of formula (11) is as follows:

[0120] ;Formula (14)

[0121] It should be pointed out that in some high-temperature thermochemical nonequilibrium flow field numerical simulations, a completely catalytic wall with the wall components given as the incoming flow condition is used. In this case, the component mass flux in formula (14) is Cannot be given explicitly.

[0122] In order to calculate the wall heat flux in the case of formula (14), the wall component mass conservation equation can be further considered:

[0123] ;Formula (15)

[0124] Formula (14) is reduced to the wall heat flux calculation formula under normal conditions:

[0125] ;Formula (16)

[0126] For active and passive ejection conditions without catalytic and ablation reactions, the wall heat flux only includes the heat conduction term of the flow field, and the calculation formula is:

[0127] ;Formula (17)

[0128] Among them, formula (17) expresses the mass fraction gradient change caused by active and passive injection into the flow field, which does not lead to the diffusion heat flow of the wall components, because this part of the energy will be balanced by the energy flux injected into the flow field. The final comprehensive effect is only the change brought about by the change of the wall temperature gradient.

[0129] Step S16: Substitute the current wall condition of the aircraft into the wall heat flux calculation equation to obtain the aerodynamic thermal analysis result of the aircraft under the corresponding wall condition.

[0130] In this embodiment, the current wall condition of the aircraft is substituted into the wall heat flux calculation equation. The wall heat flux calculation equation of the present invention can calculate the aerodynamic heat result under the current wall condition, that is, .

[0131] like Figure 4 As shown in the figure, when there is a simplified model of the aircraft as a blunt wedge shape calculation model, when the head radius R is known n =10mm, the total length of the model is 2000mm, and the injection range is 400~2000mm from the rear end of the windward surface ( Figure 4 The calculated conditions are altitude 65km, Mach number 24, angle of attack 9.5°, and wall temperature T w =1500K, under the condition.

[0132] In one embodiment, if the current wall condition is: using conventional limited catalytic conditions, only considering the catalytic composite reaction of nitrogen atoms and oxygen atoms on the wall, the catalytic coefficient is 0.01. In this case, the heat flow calculation of the present invention should adopt formula (14), and the calculation result is shown in the black curve in the figure. The traditional heat flow calculation formula still adopts formula (16), and the calculation result is shown in the red curve. Figure 5 It can be seen that the result calculated by formula (14) is consistent with the result calculated by formula (16), which confirms that the wall heat flux calculation expression of the present invention is compatible with the traditional calculation method.

[0133] In another embodiment, if the current wall condition is a non-catalytic wall with water vapor injection in the blunt wedge mass injection zone, the injection mass flow rate is 0.005 kg / (m 2 •s). In this case, the heat flow calculation of the present invention should adopt formula (17), and the traditional heat flow calculation formula still adopts formula (16).

[0134] according to Figure 6 The heat flux distribution of the blunt wedge shape wall calculated using the above two formulas, where the blue curve represents the calculated distribution of heat flux without entrainment, the red curve represents the heat flux distribution calculated by the traditional calculation formula under water vapor entrainment, and the black curve represents the heat flux distribution calculated by the formula of the present invention under water vapor entrainment. It can be seen that the heat flux obtained by the calculation formula of the present invention is lower than the heat flux without entrainment in the entire mass entrainment area, indicating that the wall gas entrainment has achieved the purpose of heat reduction, which is consistent with the actual situation, while the heat flux calculated by the traditional formula is even higher than the heat flux without entrainment near X=0.4m, and the wall gas entrainment has the effect of increasing the heat flux, which is inconsistent with the actual situation. It shows that the calculation method of the present invention can effectively evaluate the wall heat flux.

[0135] It can be seen that the present application discloses an aerodynamic thermal analysis method under multiple conditions, including: obtaining a target high-temperature gas flow field of an aircraft to extract wall parameter information and wall sublayer parameter information for evaluating wall heat flux from the convergent flow field information of the target high-temperature gas flow field; calculating the mass flux of each component flowing into and out of the control body according to the wall parameter information and the wall sublayer parameter information, and determining the conditional parameters of each wall effect condition to construct a component mass conservation equation; constructing a component mass conservation equation based on the mass flux of each component, the component diffusion conservation condition and the catalytic conservation condition in the component mass conservation equation. The total mass conservation equation of the wall; the wall energy conservation equation is constructed based on the total mass ejection flux determined by the total mass conservation equation of the wall, the wall parameter information, and the wall sub-layer parameter information; through mathematical operations on the component mass conservation equation, the total mass conservation equation of the wall, and the wall energy conservation equation, a wall heat flow calculation equation is obtained to characterize the heat flow composition of the solid structure of the aircraft under the action of each wall effect; the current wall condition of the aircraft is substituted into the wall heat flow calculation equation to obtain the aerodynamic thermal analysis results of the aircraft under the corresponding wall condition. It can be seen that by constructing the component mass conservation equation, the total mass conservation equation of the wall, and the wall energy conservation equation in detail, and obtaining the wall heat flow calculation equation based on this, the wall heat flow can be accurately calculated. This scheme comprehensively considers a variety of wall effect conditions, including wall catalysis, ablation, pyrolysis ejection, active ejection, etc., so that the wall heat flow calculation equation is applicable to wall conditions of different complexity. Whether it is a simple conventional catalytic wall without ablation or ejection, or a complex wall with multiple effects, this method can be used to perform aerodynamic thermal analysis, providing an accurate evaluation method for the thermal protection research of aircraft under various working conditions.

[0136] Reference Figure 7 As shown, the present invention also discloses a pneumatic thermal analysis device under multiple conditions, including:

[0137] An information acquisition module 11 is used to acquire a target high-temperature gas flow field of the aircraft, so as to extract wall parameter information and wall sub-layer parameter information for evaluating wall heat flux from convergent flow field information of the target high-temperature gas flow field;

[0138] A first equation building module 12 is used to calculate the mass flux of each component flowing into and out of the control volume according to the wall parameter information and the wall sub-layer parameter information, and determine the condition parameters of each wall effect condition to construct a component mass conservation equation;

[0139] A second equation building module 13 is used to build a wall total mass conservation equation based on the mass flux of each component, the component diffusion conservation condition and the catalytic conservation condition in the component mass conservation equation;

[0140] A third equation building module 14 is used to build a wall energy conservation equation based on the total mass ejection flux determined by the wall total mass conservation equation, the wall parameter information, and the wall sub-layer parameter information;

[0141] The heat flow equation determination module 15 is used to obtain a wall heat flow calculation equation that characterizes the heat flow composition felt by the solid structure of the aircraft under the action of each wall effect by performing mathematical operations on the component mass conservation equation, the wall total mass conservation equation, and the wall energy conservation equation;

[0142] The analysis module 16 is used to substitute the current wall condition of the aircraft into the wall heat flux calculation equation to obtain the aerodynamic thermal analysis result of the aircraft under the corresponding wall condition.

[0143] It can be seen that the present application discloses obtaining the target high-temperature gas flow field of the aircraft, so as to extract the wall parameter information and the wall sub-layer parameter information for evaluating the wall heat flux from the convergent flow field information of the target high-temperature gas flow field; calculating the mass flux of each component flowing into and out of the control body according to the wall parameter information and the wall sub-layer parameter information, and determining the conditional parameters of each wall effect condition to construct the component mass conservation equation; constructing the wall total mass conservation equation based on the mass flux of each component, the component diffusion conservation condition and the catalytic conservation condition in the component mass conservation equation ; Based on the total mass ejection flux determined by the wall total mass conservation equation, the wall parameter information, and the wall sub-layer parameter information, the wall energy conservation equation is constructed; through mathematical operations on the component mass conservation equation, the wall total mass conservation equation, and the wall energy conservation equation, a wall heat flux calculation equation is obtained that characterizes the heat flux composition of the solid structure of the aircraft under the action of each wall effect; the current wall condition of the aircraft is substituted into the wall heat flux calculation equation to obtain the aerodynamic thermal analysis results of the aircraft under the corresponding wall condition. It can be seen that by constructing the component mass conservation equation, the wall total mass conservation equation, and the wall energy conservation equation in detail, and obtaining the wall heat flux calculation equation based on this, the wall heat flux can be accurately calculated. This scheme comprehensively considers a variety of wall effect conditions, including wall catalysis, ablation, pyrolysis ejection, active ejection, etc., so that the wall heat flux calculation equation is applicable to wall conditions of different complexity. Whether it is a simple conventional catalytic wall without ablation or ejection, or a complex wall with multiple effects, this method can be used to perform aerodynamic thermal analysis, providing an accurate evaluation method for the thermal protection research of aircraft under various working conditions.

[0144] Furthermore, the present application also discloses an electronic device. Figure 8 This is a structural diagram of an electronic device 20 according to an exemplary embodiment. The content in the diagram cannot be regarded as any limitation on the scope of use of the present application.

[0145] Figure 8 A schematic diagram of the structure of an electronic device 20 provided in an embodiment of the present application. The electronic device 20 may specifically include: at least one processor 21, at least one memory 22, a power supply 23, a communication interface 24, an input / output interface 25, and a communication bus 26. The memory 22 is used to store a computer program, which is loaded and executed by the processor 21 to implement the relevant steps in the aerodynamic thermal analysis method under multiple conditions disclosed in any of the aforementioned embodiments. In addition, the electronic device 20 in this embodiment may specifically be an electronic computer.

[0146] In this embodiment, the power supply 23 is used to provide working voltage for each hardware device on the electronic device 20; the communication interface 24 can create a data transmission channel between the electronic device 20 and the external device, and the communication protocol it follows is any communication protocol that can be applied to the technical solution of the present application, and is not specifically limited here; the input and output interface 25 is used to obtain external input data or output data to the outside world, and its specific interface type can be selected according to specific application needs and is not specifically limited here.

[0147] Among them, the processor 21 may include one or more processing cores, such as a 4-core processor, an 8-core processor, etc. The processor 21 can be implemented in at least one hardware form of DSP (Digital Signal Processing), FPGA (Field-Programmable Gate Array), and PLA (Programmable Logic Array). The processor 21 may also include a main processor and a coprocessor. The main processor is a processor for processing data in the awake state, also known as a CPU (Central Processing Unit); the coprocessor is a low-power processor for processing data in the standby state. In some embodiments, the processor 21 may be integrated with a GPU (Graphics Processing Unit), which is responsible for rendering and drawing the content to be displayed on the display screen. In some embodiments, the processor 21 may also include an AI (Artificial Intelligence) processor, which is used to process computing operations related to machine learning.

[0148] In addition, the memory 22, as a carrier for storing resources, can be a read-only memory, a random access memory, a disk or an optical disk, etc. The resources stored thereon can include an operating system 221, a computer program 222, etc., and the storage method can be temporary storage or permanent storage.

[0149] Among them, the operating system 221 is used to manage and control the hardware devices and computer programs 222 on the electronic device 20, so as to realize the operation and processing of the massive data 223 in the memory 22 by the processor 21, which can be Windows Server, Netware, Unix, Linux, etc. In addition to including a computer program that can be used to complete the aerodynamic thermal analysis method under multiple conditions performed by the electronic device 20 disclosed in any of the aforementioned embodiments, the computer program 222 can further include a computer program that can be used to complete other specific tasks. In addition to data transmitted from an external device received by the electronic device, the data 223 can also include data collected by its own input and output interface 25, etc.

[0150] Furthermore, the present application also discloses a computer-readable storage medium for storing a computer program; wherein, when the computer program is executed by a processor, the aerodynamic thermal analysis method under multiple conditions disclosed above is implemented. For the specific steps of the method, reference may be made to the corresponding contents disclosed in the above embodiments, and no further description will be given here.

[0151] In this specification, each embodiment is described in a progressive manner, and each embodiment focuses on the differences from other embodiments. The same or similar parts between the embodiments can be referred to each other. For the device disclosed in the embodiment, since it corresponds to the method disclosed in the embodiment, the description is relatively simple, and the relevant parts can be referred to the method part.

[0152] Professionals may further appreciate that the units and algorithm steps of each example described in conjunction with the embodiments disclosed herein can be implemented with electronic hardware, computer software, or a combination of the two. In order to clearly illustrate the interchangeability of hardware and software, the composition and steps of each example have been generally described in terms of function in the above description. Whether these functions are performed in hardware or software depends on the specific application and design constraints of the technical solution. Professional and technical personnel may use different methods to implement the described functions for each specific application, but such implementation should not be considered to be beyond the scope of this application. The steps of the method or algorithm described in conjunction with the embodiments disclosed herein can be implemented directly with hardware, a software module executed by a processor, or a combination of the two. The software module can be placed in a random access memory RAM (Random Access Memory), memory, read-only memory ROM (Read Only Memory), electrically programmable EPROM (Electrically Programmable Read Only Memory), electrically erasable programmable EEPROM (ElectricErasable Programmable Read Only Memory), register, hard disk, removable disk, CD-ROM (CompactDisc-Read Only Memory), or any other form of storage medium known in the technical field.

[0153] Finally, it should be noted that, in this article, relational terms such as first and second, etc. are only used to distinguish one entity or operation from another entity or operation, and do not necessarily require or imply any such actual relationship or order between these entities or operations. Moreover, the terms "include", "comprise" or any other variants thereof are intended to cover non-exclusive inclusion, so that a process, method, article or device including a series of elements includes not only those elements, but also other elements not explicitly listed, or also includes elements inherent to such process, method, article or device. In the absence of further restrictions, the elements defined by the sentence "comprise a ..." do not exclude the presence of other identical elements in the process, method, article or device including the elements.

[0154] The scheme provided by the present invention is introduced in detail above. Specific examples are used in this article to illustrate the principle and implementation mode of the present invention. The description of the above embodiments is only used to help understand the method of the present invention and its core idea. At the same time, for those skilled in the art, according to the idea of ​​the present invention, there will be changes in the specific implementation mode and application scope. In summary, the content of this specification should not be understood as limiting the present invention.

Claims

1. A multi-condition aerodynamic thermal analysis method, characterized in that: include: Acquiring a target high-temperature gas flow field of the aircraft to extract wall parameter information and wall sub-layer parameter information for evaluating wall heat flux from convergent flow field information of the target high-temperature gas flow field; Calculating the mass flux of each component flowing into and out of the control volume according to the wall parameter information and the wall sub-layer parameter information, and determining the conditional parameters of each wall effect condition to construct a component mass conservation equation; Constructing a total wall mass conservation equation based on the mass flux of each component, the component diffusion conservation condition and the catalytic conservation condition in the component mass conservation equation; Constructing a wall energy conservation equation based on the total mass ejection flux determined by the wall total mass conservation equation, the wall parameter information, and the wall sub-layer parameter information; By performing mathematical operations on the component mass conservation equation, the wall total mass conservation equation, and the wall energy conservation equation, a wall heat flux calculation equation is obtained that characterizes the heat flux composition felt by the solid structure of the aircraft under the action of each wall effect; The current wall condition of the aircraft is substituted into the wall heat flux calculation equation to obtain the aerodynamic thermal analysis result of the aircraft under the corresponding wall condition.

2. The multi-condition aerodynamic thermal analysis method according to claim 1, characterized in that: The calculating the mass flux of each component flowing into and out of the control body according to the wall parameter information and the wall sub-layer parameter information, and determining the condition parameters of each wall effect condition, includes: Calculating the component diffusion flux of each component flowing into and out of the control body, the active ejection gas flux, the pyrolysis gas ejection flux, the catalytic reaction generation rate and the ablation reaction generation rate according to the wall parameter information and the wall sub-layer parameter information; Determine the mass proportion of each component in the actively ejected gas and the mass proportion of each component in the pyrolysis gas under each wall effect condition.

3. The multi-condition aerodynamic thermal analysis method according to claim 2, characterized in that: The wall total mass conservation equation is constructed based on the mass flux of each component in the component mass conservation equation, the component diffusion conservation condition and the catalytic conservation condition, including: The component diffusion flux, the active ejection gas flux, the pyrolysis gas ejection flux, the catalytic reaction generation rate and the ablation reaction generation rate of each component in the component mass conservation equation are added together to obtain an initial wall total mass conservation equation; Set the component diffusion conservation condition that the mass flux of each component diffusing into the control volume and the mass flux diffusing out of the control volume meet the conservation condition; A catalytic conservation condition is set in which the total mass of each reactant component and the total mass of the product component satisfy the conservation condition during the wall catalytic reaction; The total wall mass conservation equation is constructed based on the initial wall total mass conservation equation, the component diffusion conservation condition, and the catalytic conservation condition.

4. The multi-condition aerodynamic thermal analysis method according to claim 1, characterized in that: The wall energy conservation equation is constructed based on the total mass ejection flux determined by the wall total mass conservation equation, the wall parameter information, and the wall sub-layer parameter information, including: Based on the wall parameter information and the wall sub-layer parameter information, enthalpy parameters including wall component enthalpy, total wall ejected gas enthalpy, active ejected gas enthalpy, pyrolysis ejected gas enthalpy, active ejected gas generation enthalpy, pyrolysis ejected gas generation enthalpy and conduction heat flux items are calculated and determined; A wall energy conservation equation characterizing the energy transferred from the control body to the flow field and the solid structure of the aircraft is constructed according to the conduction heat flux term, the enthalpy parameter and the total mass ejection flux determined based on the wall total mass conservation equation.

5. The multi-condition aerodynamic thermal analysis method according to claim 4, characterized in that: The wall heat flow calculation equation characterizing the heat flow composition felt by the solid structure of the aircraft under the action of each wall effect is obtained by mathematically calculating the component mass conservation equation, the wall total mass conservation equation, and the wall energy conservation equation, including: Multiplying the component mass conservation equation with the wall component enthalpy value in the wall energy conservation equation and summing the values ​​for each component to obtain a total energy change equation caused by changes in each wall mass flux; According to the mathematical operation of the total energy change equation and the wall total mass conservation equation, a wall heat flux calculation equation is obtained to characterize the heat flux composition felt by the solid structure of the aircraft under the action of each wall effect.

6. The multi-condition aerodynamic thermal analysis method according to claim 5, characterized in that: The mathematical operation based on the total energy change equation and the wall total mass conservation equation is used to obtain a wall heat flow calculation equation that characterizes the heat flow composition felt by the solid structure of the aircraft under the action of each wall effect, including: According to mathematical operations of the total energy change equation and the wall total mass conservation equation, a wall heat flow calculation equation constructed based on the conduction heat flux term, the wall component enthalpy value, the active ejection gas generation enthalpy and the pyrolysis ejection gas generation enthalpy is obtained.

7. The multi-condition aerodynamic thermal analysis method according to any one of claims 1 to 6, characterized in that: The step of obtaining a target high temperature gas flow field of the aircraft comprises: The target high-temperature gas flow field of the aircraft is calculated using a preset high-temperature gas thermochemical non-equilibrium flow field numerical simulation method.

8. A multi-condition pneumatic thermal analysis device, characterized in that: include: An information acquisition module, used to acquire a target high-temperature gas flow field of the aircraft, so as to extract wall parameter information and wall sub-layer parameter information for evaluating wall heat flux from convergent flow field information of the target high-temperature gas flow field; A first equation building module is used to calculate the mass flux of each component flowing into and out of the control volume according to the wall parameter information and the wall sub-layer parameter information, and determine the condition parameters of each wall effect condition to construct a component mass conservation equation; A second equation building module is used to build a wall total mass conservation equation based on the mass flux of each component, the component diffusion conservation condition and the catalytic conservation condition in the component mass conservation equation; A third equation building module is used to build a wall energy conservation equation based on the total mass ejection flux determined by the wall total mass conservation equation, the wall parameter information, and the wall sub-layer parameter information; A heat flow equation determination module is used to obtain a wall heat flow calculation equation that characterizes the heat flow composition felt by the solid structure of the aircraft under the action of each wall effect by performing mathematical operations on the component mass conservation equation, the wall total mass conservation equation, and the wall energy conservation equation; The analysis module is used to substitute the current wall condition of the aircraft into the wall heat flux calculation equation to obtain the aerodynamic thermal analysis result of the aircraft under the corresponding wall condition.

9. An electronic device, characterized in that: include: Memory, used to store computer programs; A processor is used to execute the computer program to implement the steps of the aerodynamic thermal analysis method under multiple conditions as described in any one of claims 1 to 7.

10. A computer-readable storage medium, characterized in that: Used to store computer programs; wherein, when the computer program is executed by a processor, the steps of the aerodynamic thermal analysis method under multiple conditions as described in any one of claims 1 to 7 are implemented.

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