Method for calculating coupling thermal environment of hypersonic flight vehicle along flight envelope

Abstract

The invention discloses a method for calculating a coupling thermal environment of a hypersonic flight vehicle along a flight envelope, and belongs to the technical field of hypersonic flight vehicle design. The problem of thermal environment calculation of long-time dynamic flight along a flight envelope is solved. The method comprises the following steps: constructing an aerodynamic heat calculation model; constructing a structural thermal environment calculation model for calculating coupling wall surface temperature data; constructing a thermal radiation calculation correction model for calculating the influence quantity of outward thermal radiation on the heat flow; constructing a coupling calculation model for realizing coupling boundary data transmission; on the basis of the models, a coupling boundary data transmission method is built, next-moment propulsion calculation of a flow field and a structure field is carried out, an internal pseudo-time iteration mode is adopted till the iteration convergence requirement is met, and then a hypersonic speed coupling thermal environment data calculation result is obtained. According to the invention, the tight coupling and accurate data transmission process of multiple physical fields can be realized, and the problems of coupling thermal environment, thermal management, thermal protection and accurate prediction of thermodynamic elasticity of the hypersonic aircraft are solved.

Description

Technical Field

[0001] The present invention relates to the field of aerodynamic heating, thermal radiation and structural thermal environment prediction for the design and development of hypersonic aircraft. Specifically, it provides a method for calculating the coupled thermal environment of a hypersonic aircraft along its flight envelope. The method establishes a high-precision numerical calculation model for multi-physics field coupled prediction of aerodynamic heating / structural heat conduction / thermal radiation, meeting the demand for rapid and accurate prediction of the thermal environment during the aircraft design and development stage. Background Art

[0002] Near-space hypersonic vehicles have attracted widespread attention from major aerospace powers worldwide due to their significant technical advantages and practical value in flight environments with wide speed ranges and large airspaces. During long, cross-domain, high-speed, dynamic flight, intense friction between the vehicle and the air generates aerodynamic heating. This results in prolonged high-temperature exposure, impacting the aircraft's skin, structure, thermal protection, infrared stealth, and thermal management. Accurately predicting aerodynamic heating has become a critical technical challenge in aircraft development.

[0003] In engineering, numerical calculations or high-enthalpy wind tunnel aerodynamic thermal tests are usually used to obtain the heat flux or temperature distribution on the aircraft surface. The boundary conditions of the numerical simulation are usually set as isothermal walls or adiabatic walls, and the structural thermal environment or structural strength is evaluated based on this as the thermal boundary data. Due to the large range of changes in flight altitude and flight speed during long-term dynamic flight, and the long-term thermal radiation generated by the high temperature environment on the aircraft surface, the above multiple factors will have a significant impact on the thermal environment of the aircraft structure, resulting in excessive redundancy in the aircraft's thermal protection and thermal design, which restricts the overall performance improvement of the aircraft.

[0004] The present invention is a hypersonic aircraft aerodynamic / structural / radiation coupled thermal environment prediction model and data transmission method, which constructs a thermal environment prediction model of aerodynamic heating, structural heat conduction, and thermal radiation coupling, a tight coupling algorithm and a data transmission method, and solves the cumulative effect of thermal radiation during long-term dynamic flight. Summary of the Invention

[0005] In order to meet the requirements of refined design and rapid iteration of modern near-space hypersonic aircraft, and to address the shortcomings of traditional CFD aerodynamic thermal calculations that cannot meet the requirements of refined aircraft design, the present invention proposes a coupled thermal environment calculation method for hypersonic aircraft along the flight envelope, aiming to solve the problems of thermal environment and radiation cumulative effects during long-term dynamic flight of hypersonic aircraft.

[0006] To achieve the above object, the present invention is implemented through the following technical solutions:

[0007] A method for calculating the coupled thermal environment of a hypersonic vehicle along its flight envelope comprises the following steps:

[0008] S1. Construct an aerodynamic heat calculation model to calculate the heat flux distribution data on the aircraft surface under isothermal wall conditions;

[0009] S2. Construct a structural thermal environment calculation model to calculate the structural temperature field distribution data;

[0010] S3. Construct a thermal radiation calculation correction model to calculate the effect of outward thermal radiation on heat flux after the temperature of the aircraft wall increases;

[0011] S4. Construct a coupled computational model to implement coupled boundary heat flow and temperature data transfer;

[0012] S5. Based on the model from steps S1-S4, a coupled boundary data transfer method is constructed. For the coupled boundary, the flow field sends the calculated heat flux value to the structural field. After receiving the flow field heat flux value, the structural field uses a thermal radiation calculation correction model to subtract the effect of thermal radiation on aerodynamic heating. Then, the structural temperature field is calculated based on the structural thermal environment calculation model to obtain the coupled boundary temperature value. Furthermore, the structural field sends the coupled boundary temperature value to the flow field. Upon receiving the coupled boundary temperature value, the flow field assigns the temperature value to the flow field coupled boundary as the flow field temperature boundary condition. This completes the coupled boundary data transfer.

[0013] Then, the next moment advance calculation of the flow field and structural field is carried out, and the internal pseudo-time iteration method is used in the advance calculation until the iterative convergence requirements are met. The calculation results of the hypersonic coupled thermal environment data at the next moment are obtained, and the calculation is carried out until all the calculation contents are completed.

[0014] Furthermore, step S1 is to calculate the aerodynamic and thermal characteristics of the entire aircraft. The CFD method is used to obtain the information of the entire aircraft and its walls, including lift, drag, pressure, heat flow, temperature and density parameters. Without considering the body force and external heat source, the aerodynamic and thermal calculation model is constructed as follows:

[0015]

[0016] Where Q is the conservation flux of the NS equation, 、 and are the inviscid vector fluxes of the convection term in the x, y and z directions respectively, 、 and are the viscous vector fluxes in the x, y and z directions respectively.

[0017] Furthermore, the structural thermal environment calculation model in step S2 utilizes the method of structural computational heat transfer to realize the dynamic calculation of the non-uniformly distributed temperature field of the aircraft skin and internal structure. The unsteady-state heat conduction equation under the three-dimensional isotropic assumption is as follows:

[0018]

[0019] in, is the temperature, is the solid density, is the specific heat of the solid at constant pressure, is the thermal conductivity of the solid;

[0020] Furthermore, in step S3, a thermal radiation calculation model is constructed to consider the impact of the thermal radiation effect between the space and the high-temperature wall on the structural thermal environment during the long-term dynamic flight of the aircraft. The blackbody radiation Stefan-Boltzmann law is used to quantitatively describe the radiation per unit time per unit area. The thermal radiation effect correction amount is calculated based on the aircraft surface temperature distribution data obtained in step S2:

[0021]

[0022] in, Indicates the thermal radiation influence data, represents the emissivity of the object, represents the Stefan-Boltzmann constant, which is , and represent the wall temperature and the incoming flow temperature respectively.

[0023] Furthermore, in step S4, a coupling calculation model is constructed to obtain:

[0024]

[0025] in, is the heat flux value under isothermal wall conditions obtained by numerical calculation of the flow field, is the heat flux value that conducts heat into the structure through the coupled boundary.

[0026] Furthermore, the specific implementation method of step S5 includes the following steps:

[0027] S5.1. Discretize the aircraft into a set of computational state points based on its specific flight envelope or state. Flight states include time, Mach number, angle of attack, pressure, density, temperature, etc.

[0028] S5.2. Perform steady-state calculations of the aircraft flow field based on the initial state of the flight state point set to obtain the flow field in this state. This flow field serves as the initial flow field for the unsteady calculations.

[0029] S5.3. Perform coupled boundary data transfer calculations;

[0030] S5.3.1. Extract the initial flow field (at this moment) of the unsteady flow calculation and calculate the wall heat flux.

[0031] S5.3.2. Transfer the current flow field coupled boundary heat flux data to the structural field.

[0032] S5.3.3. After receiving the coupled boundary heat flux value, the structural field subtracts the thermal radiation effect at the current temperature value and performs structural heat transfer calculations to obtain the structural temperature distribution data at the next moment.

[0033] S5.3.4. The structural field transmits the coupled boundary temperature data at the next moment to the flow field, assigns it to the flow field coupled boundary, and updates the coupled boundary temperature data.

[0034] S5.3.5. Simultaneously perform unsteady advance calculations of the flow field aerodynamic heat transfer and structural heat conduction. Use an internal pseudo-time iteration method until convergence is achieved. Then, assign the internal iteration data to the next time step.

[0035] S5.4. Use the iterative convergence data of the flow and structural fields as the results for the next moment;

[0036] S5.5. Repeat S5.3-S5.4 until all the state point sets are calculated.

[0037] Beneficial effects of the present invention:

[0038] The method described in this paper calculates the coupled thermal environment of a hypersonic aircraft along its flight envelope. It constructs a coupled thermal environment prediction model for aerodynamic heating, structural heat conduction, and thermal radiation, along with a tightly coupled algorithm and data transfer method. This method addresses the cumulative effects of thermal radiation during long-term dynamic flight. It achieves tight coupling of multiple physical fields and precise data transfer, addressing the challenges of accurately predicting the coupled thermal environment, thermal management, thermal protection, and thermoaeroelasticity of hypersonic aircraft. BRIEF DESCRIPTION OF THE DRAWINGS

[0039] Figure 1 This is a flow chart of a method for calculating the coupled thermal environment of a hypersonic aircraft along its flight envelope, as proposed by the present invention;

[0040] Figure 2 Schematic diagram of the cylindrical model and mesh;

[0041] Figure 3 The flow field, structural field and coupled boundary temperature distribution of the cylindrical model at flight altitude H = 25 km, Ma = 4, and flight time t = 100 s;

[0042] Figure 4 The wind tunnel test model of a two-dimensional wing with trusses and the schematic diagram of the grid distribution;

[0043] Figure 5 The upper surface temperature distribution of the wind tunnel test model of a two-dimensional wing with trusses at test time t=5s;

[0044] Figure 6 The temperature distribution on the lower surface of the wind tunnel test model of a two-dimensional wing with trusses at test time t=5s;

[0045] Figure 7 The temperature distribution of the cross-section of the wind tunnel test model with trusses at test time t=5s;

[0046] Figure 8 The temperature distribution of the coupled boundary section of the wind tunnel test model of a two-dimensional wing with trusses at test time t=5s. DETAILED DESCRIPTION

[0047] In order to make the objectives, technical solutions, and advantages of the present invention more clearly understood, the present invention is further described in detail below with reference to the accompanying drawings and specific embodiments. It should be understood that the specific embodiments described herein are only intended to explain the present invention and are not intended to limit the present invention. That is, the specific embodiments described herein are only some embodiments of the present invention, not all embodiments. Generally, the components of the specific embodiments of the present invention described and illustrated in the drawings herein can be arranged and designed in various different configurations, and the present invention can also have other embodiments.

[0048] Therefore, the following detailed description of the specific embodiments of the present invention provided in the accompanying drawings is not intended to limit the scope of the invention as claimed, but is merely representative of selected specific embodiments of the present invention. All other specific embodiments obtained by those skilled in the art based on the specific embodiments of the present invention without making any creative efforts shall fall within the scope of protection of the present invention.

[0049] In order to further understand the content, features and effects of the present invention, the following specific embodiments are given as examples, and the attached Figure 1 -Attached Figure 8 The detailed instructions are as follows:

[0050] Example 1:

[0051] A method for calculating hypersonic coupled thermal environment data comprises the following steps:

[0052] S1. Construct an aerodynamic thermal calculation model to calculate the heat flux distribution data on the aircraft surface;

[0053] Furthermore, step S1 is to calculate the aerodynamic and thermal characteristics of the entire aircraft. The CFD method is used to obtain the information of the entire aircraft and its walls, including lift, drag, pressure, heat flow, temperature and density parameters. Without considering the body force and external heat source, the aerodynamic and thermal calculation model is constructed as follows:

[0054]

[0055] Where Q is the conservation flux of the NS (Navier-Stokes) equation, 、 and are the inviscid vector fluxes of the convection term in the x, y and z directions respectively, 、 and are the viscous vector fluxes in the x, y and z directions respectively.

[0056] Further,

[0057]

[0058]

[0059]

[0060] in, represents the viscous stress term, are the velocity components in the Cartesian coordinate system, is the total energy per unit mass of gas, is the fluid density, is the fluid pressure;

[0061] The specific form is:

[0062]

[0063] in, is the heat flow, is the dynamic viscosity coefficient;

[0064] The specific formula is as follows:

[0065]

[0066] in, is the specific heat ratio;

[0067]

[0068] in, It represents the internal energy per unit mass of gas;

[0069] S2. Construct a structural thermal environment calculation model to calculate structural temperature distribution data;

[0070] Furthermore, the structural thermal environment calculation model in step S2 utilizes the method of structural computational heat transfer to realize the dynamic calculation of the non-uniformly distributed temperature field of the aircraft skin and internal structure. The unsteady-state heat conduction equation under the three-dimensional isotropic assumption is as follows:

[0071]

[0072] in, is the temperature, is the solid density, is the specific heat of the solid at constant pressure, is the thermal conductivity of the solid;

[0073] Taking the curvilinear coordinate system equation of the two-dimensional dimensionless heat conduction equation as an example, its spatial and temporal discretization forms are introduced:

[0074]

[0075] in, The Jacobian of the coordinate transformation is represented by is the first curvilinear coordinate;

[0076] ; ; ;

[0077] in, is the second curve coordinate;

[0078] , ,

[0079] In the time advancement process of the unsteady calculation of structural heat conduction, the time derivative is discretized using the implicit three-point back-difference method with second-order time accuracy, and the unsteady implicit heat conduction discrete equation with second-order time accuracy is obtained:

[0080]

[0081] in, Represents the coordinate point position under the curve coordinates, represents the real time step, and R represents the right-hand side term;

[0082] In order to ensure the second-order accuracy of the unsteady calculation, the time derivative term is solved by the dual time step method, a virtual time derivative term is added to the time discrete term, and a first-order forward difference is used to introduce a pseudo time step. , we get the numerical method of the unsteady heat conduction equation:

[0083]

[0084] in, and are the virtual time step and the real time step, p and n are the number of virtual time iteration steps and the number of real time advancement steps, respectively; is the temperature value at time p, The right-hand term.

[0085] S3. Construct a thermal radiation calculation correction model to calculate thermal radiation impact data;

[0086] Furthermore, in step S3, a thermal radiation calculation model is constructed to consider the impact of the thermal radiation effect between the space and the high-temperature wall on the structural thermal environment during the long-term dynamic flight of the aircraft. The blackbody radiation Stefan-Boltzmann law is used to quantitatively describe the radiation per unit time per unit area. The thermal radiation effect correction amount is calculated based on the aircraft surface temperature distribution data obtained in step S2:

[0087]

[0088] in, Indicates the thermal radiation influence data, represents the emissivity of the object, represents the Stefan-Boltzmann constant, which is , and represent the wall temperature and the incoming flow temperature respectively.

[0089] S4. Construct a coupled computational model to implement coupled boundary data transfer;

[0090] Furthermore, in step S4, a coupling calculation model is constructed to obtain:

[0091]

[0092] in, is the heat flux value obtained by numerical calculation of the flow field, is the coupled boundary heat flux value for heat conduction into the structure.

[0093] S5. Based on the model established in steps S1-S4, a coupled boundary data transfer method is constructed. For the coupled boundary, the flow field sends the calculated heat flux value to the structural field. After receiving the flow field heat flux value, the structural field uses a thermal radiation calculation correction model to subtract the thermal radiation influence data. Simultaneously, the structural temperature field is calculated based on the structural thermal environment calculation model to obtain the coupled boundary temperature value of the structural field. Furthermore, the structural field sends the coupled boundary temperature value to the flow field. After receiving the coupled boundary temperature value, the flow field assigns the temperature value to the flow field coupled boundary as the flow field temperature boundary condition. This completes the coupled boundary data transfer.

[0094] Then, the next moment advance calculation of the flow field and structural field is carried out, and the internal pseudo-time iteration method is used until the convergence requirements are met, and the calculation results of the hypersonic coupled thermal environment data are obtained.

[0095] Furthermore, the specific implementation method of step S5 includes the following steps:

[0096] S5.1. Discretize the aircraft into a set of computational state points based on its specific flight envelope or state. Flight states include time, Mach number, angle of attack, pressure, density, temperature, etc.

[0097] S5.2. Perform steady-state calculations of the aircraft flow field based on the initial state of the flight state point set to obtain the flow field in this state. This flow field serves as the initial flow field for the unsteady calculations.

[0098] S5.3. Perform coupled boundary data transfer calculations;

[0099] S5.3.1. Extract the initial flow field (at this moment) of the unsteady flow calculation and calculate the wall heat flux.

[0100] S5.3.2. Transfer the current flow field coupled boundary heat flux data to the structural field.

[0101] S5.3.3. After receiving the coupled boundary heat flux value, the structural field subtracts the thermal radiation effect at the current temperature value and performs structural heat transfer calculations to obtain the structural temperature distribution data at the next moment.

[0102] S5.3.4. The structural field transmits the coupled boundary temperature data at the next moment to the flow field, assigns it to the flow field coupled boundary, and updates the coupled boundary temperature data.

[0103] S5.3.5. Simultaneously perform unsteady advance calculations of the flow field aerodynamic heat transfer and structural heat conduction. Use an internal pseudo-time iteration method until convergence is achieved. Then, assign the internal iteration data to the next time step.

[0104] S5.4. Use the iterative convergence data of the flow and structural fields as the results for the next moment;

[0105] S5.5. Repeat S5.3-S5.4 until all the state point sets are calculated.

[0106] Figure 2 It is a cylindrical model and grid diagram. The flow field and structural field are point-to-point connected and divided by the coupling boundary. That is, the flow field carries out aerodynamic heat numerical calculation, and the structural field carries out temperature heat transfer numerical calculation. Data transfer is completed at the coupling boundary.

[0107] Figure 3 The following are the flow field, structural field, and coupled boundary temperature distribution diagrams for a cylindrical model at a flight altitude of H = 25 km, Ma = 4, and a flight time of t = 100 seconds. Within 100 seconds, thermal radiation affects the surface temperature of the cylindrical model by 10 degrees. Therefore, the cumulative effect of infrared radiation from a long-duration cruise aircraft is significant. Constructing a coupled thermal environment calculation method for an aircraft, combining aerodynamic heat, structural heat conduction, and thermal radiation, can more accurately predict the thermal environment along the flight envelope.

[0108] Figure 4 The wind tunnel test model of a two-dimensional wing with trusses and the schematic diagram of the grid distribution show that the wind tunnel test model of a two-dimensional wing with trusses consists of a skin and a truss structure with a cavity in the middle and the coupling boundary is the outer wall of the skin.

[0109] Figure 5 and Figure 6 They are the temperature distribution diagrams of the upper and lower surfaces of the wind tunnel test model with a truss two-dimensional wing at the test time t=5s. The temperature value is the data transferred from the structure field to the flow field;

[0110] Figure 7 The temperature distribution diagram of the cross-section of the wind tunnel test model with a truss 2D wing at test time t=5s is obtained through structural heat conduction calculation. The outer wall boundary of the skin is the aerodynamic thermal data of the flow field.

[0111] Figure 8 The temperature distribution diagram of the coupled boundary section of the wind tunnel test model of a two-dimensional wing with a truss at test time t=5s shows that the temperature distribution varies due to different structures at different positions. It is necessary to establish an accurate structural heat conduction model to accurately reproduce the topology of the structural model.

[0112] It should be noted that relational terms such as "first" and "second" are used only to distinguish one entity or operation from another entity or operation, and do not necessarily require or imply any actual relationship or order between these entities or operations. Moreover, the terms "comprises," "comprising," or any other variations thereof are intended to cover non-exclusive inclusion, such that a process, method, article, or apparatus comprising a series of elements includes not only those elements, but also other elements not explicitly listed, or elements inherent to such process, method, article, or apparatus. In the absence of further limitations, an element defined by the phrase "comprising a ..." does not exclude the presence of additional identical elements in the process, method, article, or apparatus comprising the element.

[0113] Although the present application has been described above with reference to specific embodiments, various modifications may be made thereto and components may be substituted with equivalents without departing from the scope of the present application. In particular, as long as there are no structural conflicts, the various features of the embodiments disclosed herein may be combined with each other in any manner, and the omission of an exhaustive description of these combinations in this specification is solely for the sake of space and resource conservation. Therefore, the present application is not limited to the specific embodiments disclosed herein, but includes all technical solutions within the scope of the claims.

Claims

1. A method for calculating the coupled thermal environment of a hypersonic vehicle along its flight envelope, characterized in that: The steps include: S1. Construct an aerodynamic heat calculation model to calculate the heat flux distribution data on the aircraft surface under isothermal wall conditions; S2. Construct a structural thermal environment calculation model to calculate the structural temperature field distribution data; S3. Construct a thermal radiation calculation correction model to calculate the effect of outward thermal radiation on heat flux after the temperature of the aircraft wall increases; S4. Construct a coupled computational model to implement coupled boundary heat flow and temperature data transfer; S5. Based on the model from steps S1-S4, a coupled boundary data transfer method is constructed. For the coupled boundary, the flow field sends the calculated heat flux value to the structural field. After receiving the flow field heat flux value, the structural field uses a thermal radiation calculation correction model to subtract the effect of thermal radiation on aerodynamic heating. Then, the structural temperature field is calculated based on the structural thermal environment calculation model to obtain the coupled boundary temperature value. Furthermore, the structural field sends the coupled boundary temperature value to the flow field. Upon receiving the coupled boundary temperature value, the flow field assigns the temperature value to the flow field coupled boundary as the flow field temperature boundary condition. This completes the coupled boundary data transfer. Then, the next moment advance calculation of the flow field and structural field is carried out, and the internal pseudo-time iteration method is used in the advance calculation until the iterative convergence requirements are met. The calculation results of the hypersonic coupled thermal environment data at the next moment are obtained, and the calculation is carried out until all the calculation contents are completed.

2. The method for calculating the coupled thermal environment of a hypersonic vehicle along its flight envelope according to claim 1, characterized in that: Step S1 is to calculate the aerodynamic and thermal characteristics of the entire aircraft. The CFD method is used to obtain the aircraft and wall information, including lift, drag, pressure, heat flow, temperature and density parameters. Without considering the body force and external heat source, the aerodynamic and thermal calculation model is constructed as follows: ； Where Q is the conservation flux of the NS equation, 、 and are the inviscid vector fluxes of the convection term in the x, y and z directions respectively, 、 and are the viscous vector fluxes in the x, y and z directions respectively.

3. A method for calculating the coupled thermal environment of a hypersonic vehicle along its flight envelope according to claim 1 or 2, characterized in that: The structural thermal environment calculation model in step S2 uses the method of structural computational heat transfer to realize the dynamic calculation of the non-uniform temperature field of the aircraft skin and internal structure. The unsteady heat conduction equation under the three-dimensional isotropic assumption is as follows: ； in, is the temperature, is the solid density, is the specific heat of the solid at constant pressure, is the thermal conductivity of the solid.

4. The method for calculating the coupled thermal environment of a hypersonic vehicle along its flight envelope according to claim 3, wherein: In step S3, a thermal radiation calculation model is constructed to consider the impact of the thermal radiation effect between space and high-temperature walls on the structural thermal environment during the long-term dynamic flight of the aircraft. The blackbody radiation Stefan-Boltzmann law is used to quantitatively describe the radiation per unit time per unit area. The thermal radiation effect correction value is calculated based on the aircraft surface temperature distribution data obtained in step S2: ； in, Indicates the thermal radiation influence data, represents the emissivity of the object, represents the Stefan-Boltzmann constant, which is , and represent the wall temperature and the incoming flow temperature respectively.

5. The method for calculating the coupled thermal environment of a hypersonic vehicle along its flight envelope according to claim 4, characterized in that: The coupling calculation model constructed in step S4 is: ； in, is the heat flux value under isothermal wall conditions obtained by numerical calculation of the flow field, is the heat flux value that conducts heat into the structure through the coupled boundary.

6. The method for calculating the coupled thermal environment of a hypersonic vehicle along its flight envelope according to claim 5, characterized in that: The specific implementation method of step S5 includes the following steps: S5.

1. Discretize the aircraft into a set of computational state points based on its specific flight envelope or state. Flight states include time, Mach number, angle of attack, pressure, density, temperature, etc. S5.

2. Perform steady-state calculations of the aircraft flow field based on the initial state of the flight state point set to obtain the flow field in this state. This flow field serves as the initial flow field for the unsteady calculations. S5.

3. Perform coupled boundary data transfer calculations; S5.3.

1. Extract the initial flow field (at this moment) of the unsteady flow calculation and calculate the wall heat flux. S5.3.

2. Transfer the current flow field coupled boundary heat flux data to the structural field. S5.3.

3. After receiving the coupled boundary heat flux value, the structural field subtracts the thermal radiation effect at the current temperature value and performs structural heat transfer calculations to obtain the structural temperature distribution data at the next moment. S5.3.

4. The structural field transmits the coupled boundary temperature data at the next moment to the flow field, assigns it to the flow field coupled boundary, and updates the coupled boundary temperature data. S5.3.

5. Simultaneously perform unsteady advance calculations of the flow field aerodynamic heat transfer and structural heat conduction. Use an internal pseudo-time iteration method until convergence is achieved. Then, assign the internal iteration data to the next time step. S5.

4. Use the iterative convergence data of the flow and structural fields as the results for the next moment; S5.

5. Repeat S5.3-S5.4 until all the state point sets are calculated.

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