Method for rapidly predicting thermal environment of aircraft skin structure along flight envelope

Abstract

The invention discloses a method for rapidly predicting a thermal environment of an aircraft skin structure along a flight envelope, and belongs to the technical field of structural thermal environment prediction. The problems that in the prior art, thermal environment distribution data obtained through a traditional aircraft skin structure thermal environment rapid prediction method is low in precision and low in calculation efficiency are solved. The method comprises the following steps: discretizing a flight envelope to form a flight envelope point set, carrying out aerodynamic heat environment prediction, and carrying out aerodynamic heat data numerical simulation by utilizing a CFD numerical calculation method; according to different wall surface temperatures and heat insulation wall temperatures, a heat flow wall surface temperature correction model is constructed; thermal environment simulation of the aircraft skin structure along the flight envelope is carried out, the extracted temperature of the aircraft skin structure is corrected through a heat flow wall surface temperature correction model, propulsion calculation along the flight envelope time dimension is circularly carried out, and rapid prediction of the thermal environment is achieved. The method effectively improves the accuracy of prediction data, and can be applied to the thermal environment prediction of the hypersonic aircraft structure.

Description

Technical Field

[0001] The invention relates to a method for quickly predicting the thermal environment of an aircraft skin structure, and in particular to a method for quickly predicting the thermal environment of an aircraft skin structure along a flight envelope, and belongs to the technical field of structural thermal environment prediction. Background Art

[0002] Near-space hypersonic vehicles have attracted widespread attention due to their huge technical advantages and use value in a wide speed range and large airspace flight environment. During long-term cross-domain high-speed dynamic flight, the aircraft suffers intense friction with the air and produces aerodynamic heating, which causes the aircraft to be in a high-temperature environment for a long time during high-speed flight, producing a series of impacts on the skin, structure, thermal protection, infrared stealth and thermal management. The accurate prediction of the temperature field of the heated structure of the aircraft skin caused by aerodynamic heating has become the most urgent technical problem that needs to be solved in the development of aircraft.

[0003] With the development of aviation technology, aircraft design is moving towards precision, refinement and rapid iteration. The aerodynamic shape design of an aircraft no longer only considers the aerodynamic characteristics themselves. Aerodynamic thermal design has become the key to the success of research and development. The thermal environment is generated by the intense friction between the aircraft surface skin and the air, and acts on the skin structure to be transmitted internally. It involves comprehensive predictions of multidisciplinary and multi-physical fields such as aerodynamics, structure, and thermal management. This puts higher requirements on the high-precision prediction of the aerodynamic or structural coupled thermal environment in the design and development of hypersonic aircraft, and requires facing more complex problems and solutions than before.

[0004] In engineering, the aerodynamic thermal environment and the structural thermal environment are usually considered separately. First, the temperature or heat flux distribution under a single-point flight state in the flight envelope is obtained using isothermal wall or adiabatic wall conditions. The structural thermal environment uses the heat flux or temperature distribution of the aerodynamic thermal environment under typical working conditions or typical heating conditions as input to predict the structural temperature field. Usually only one iteration is performed, ignoring the dynamic changes in the aerodynamic thermal environment caused by changes in flight conditions during dynamic flight of the aircraft and different heat distribution on the aircraft skin surface. The structural thermal environment prediction also selects the most severe point in the flight process as input for assessment. The calculation accuracy often cannot meet the requirements for high data accuracy in aircraft design, and the resulting design margin is often too large. The calculation cost and efficiency generated by multiple rounds of iterative calculations cannot meet the design requirements.

[0005] In summary, there is a need for a method for rapid prediction of the thermal environment of an aircraft skin structure that takes into account the effect of wall temperature changes on heat flux, has high computational efficiency, and obtains accurate results. Summary of the invention

[0006] A brief overview of the present invention is provided below in order to provide a basic understanding of certain aspects of the present invention. It should be understood that this overview is not an exhaustive overview of the present invention. It is not intended to identify key or important parts of the present invention, nor is it intended to limit the scope of the present invention. Its purpose is merely to present certain concepts in a simplified form as a prelude to a more detailed description discussed later.

[0007] In view of this, in order to solve the problem that the thermal environment distribution data obtained by the traditional aircraft skin structure thermal environment rapid prediction method in the prior art has low accuracy and slow calculation efficiency, the present invention provides an aircraft skin structure thermal environment rapid prediction method.

[0008] The technical solution is as follows: A method for quickly predicting the thermal environment of an aircraft skin structure along a flight envelope comprises the following steps:

[0009] S1. Discretize the flight envelope to form a flight envelope point set for subsequent propulsion calculation along the time dimension of the flight envelope, that is, use discontinuity points to describe the dynamic flight process of the aircraft along the flight envelope;

[0010] Specifically: the flight state of each state point of the flight envelope point set includes flight time t, altitude H, Mach number Ma, angle of attack α, fluid pressure p, fluid density ρ and wall temperature T;

[0011] S2. Carry out aerodynamic thermal environment prediction and use CFD numerical calculation method to perform numerical simulation of aerodynamic thermal data under the flight state of the flight envelope point set;

[0012] Specifically: Aerothermal data include wall heat flux under isothermal wall boundary conditions and adiabatic wall temperature under adiabatic wall boundary conditions;

[0013] S3. constructing a correction model of heat flux affected by wall temperature by establishing a linear or quasi-linear temperature relationship according to heat flux values ​​under isothermal wall boundary conditions with different wall temperatures and adiabatic wall temperatures under adiabatic wall boundary conditions;

[0014] S4. Carry out thermal environment simulation of the aircraft skin structure along the flight envelope, correct the extracted aircraft skin structure temperature through the correction model of heat flux affected by wall temperature, and repeat the advancement calculation along the time dimension of the flight envelope until the thermal environment data of the skin structure at all state points under the flight envelope are obtained to realize rapid prediction of the thermal environment.

[0015] Further, in S2, aerodynamic thermal numerical simulation is performed by solving the three-dimensional unsteady compressible NS equations, and the body force and external heat source are not considered in the solution process, and the three-dimensional unsteady compressible NS equations are dimensionless;

[0016] The dimensionless form of the differential form of the three-dimensional unsteady compressible NS equations with respect to the conserved variable Q in the Cartesian coordinate system is expressed as:

[0017]

[0018] Where x is the horizontal axis of the Cartesian coordinate system, y is the vertical axis of the Cartesian coordinate system, z is the vertical axis of the Cartesian coordinate system, F is the inviscid vector flux of the first convection term, G is the inviscid vector flux of the second convection term, H is the inviscid vector flux of the third convection term, and F ν is the first viscous flux vector, G ν is the second viscous flux vector, H ν is the third viscous flux vector;

[0019] The conserved variable Q is expressed as:

[0020]

[0021] Wherein, u is the first velocity component in the Cartesian coordinate system, v is the second velocity component in the Cartesian coordinate system, w is the third velocity component in the Cartesian coordinate system, and E is the total energy per unit mass of gas;

[0022] The first convection term, the inviscid flux vector F, is expressed as:

[0023]

[0024] The second convection term, the inviscid flux vector G, is expressed as:

[0025]

[0026] The third convection term, the inviscid vector flux H, is expressed as:

[0027]

[0028] First viscous flux vector F ν It is expressed as:

[0029]

[0030]

[0031]

[0032]

[0033] Among them, τ xx is the first viscous stress term of the first viscous vector flux, τ xy is the second viscous stress term of the first viscous vector flux, τ xz is the third viscous stress term of the first viscous flux vector, is the fourth viscous stress term of the first viscous flux vector, q x is the heat flow in the x direction;

[0034] The second viscous flux G ν It is expressed as:

[0035]

[0036]

[0037]

[0038]

[0039] Among them, τ yx is the first viscous stress term of the second viscous vector flux, τ yy is the second viscous stress term of the second viscous vector flux, τ yz is the third viscous stress term of the second viscous flux vector, is the fourth viscous stress term of the second viscous vector flux, q y is the heat flow in the y direction;

[0040] The third viscous flux vector H ν It is expressed as:

[0041]

[0042]

[0043]

[0044]

[0045] Among them, τ zx is the first viscous stress term of the third viscous vector flux, τ zy is the second viscous stress term of the third viscous vector flux, τ zz is the third viscous stress term of the third viscous vector flux, is the fourth viscous stress term of the third viscous vector flux, q z is the heat flow in z direction;

[0046] The total energy E of a unit mass of gas is expressed as:

[0047]

[0048]

[0049] Among them, e is the internal energy per unit mass of gas, γ is the specific heat ratio;

[0050] Wall heat flux q w It is expressed as:

[0051]

[0052]

[0053] Where k is the convective heat transfer coefficient, μ is the molecular viscosity coefficient, c p is the specific heat capacity at constant pressure, r is the gas constant, is the normal unit length outside the wall, and the partial derivative of temperature and the normal unit length outside the wall By T x , T y and T z Solve, T x , T y and T z represent the temperature gradients in the x, y, and z directions, respectively;

[0054] T x , T y and T z Respectively expressed as:

[0055] T x =T ξ ξ x +T η η x +T ζ ζ x

[0056] T y =T ξ ξ y +T η η y +T ζ ζ y

[0057] T z =T ξ ξ z +T η η z +T ζ ζ z

[0058] Among them, ξ i , η i and i All represent coordinate conversion coefficients, i = x, y, z, T ξ , T η and T ζ They represent the temperature gradients in the directions of ξ, η and ζ respectively.

[0059] Furthermore, in S3, the correction model of heat flux affected by wall temperature is expressed as:

[0060]

[0061] Among them, q w (T_w) is the heat flux value under isothermal wall boundary conditions when the wall temperature is T_w, T_ad is the adiabatic wall temperature, q w (T_P) is the heat flux value when the wall temperature is T_P.

[0062] Furthermore, in the S4, the heat flux value of the first state point of the flight envelope point set is passed as input to the heat conduction calculation program, and the advancement calculation is carried out along the time dimension of the flight envelope until the calculation is completed and reaches the second state point. According to the extracted skin structure surface temperature distribution data, the heat flux value of each point of the aircraft surface grid is corrected and calculated through a correction model of heat flux affected by wall temperature to obtain the corrected heat flux value of the second state point, and the corrected heat flux data is passed to the heat conduction calculation program, and the advancement calculation is carried out along the time dimension of the flight envelope until the calculation is completed and reaches the next state point. The above steps are repeated, and the solution is successively advanced until the calculation of all state points is completed, and the thermal environment distribution of the aircraft skin structure along the flight envelope is obtained.

[0063] The beneficial effects of the present invention are as follows: the present invention integrates aerodynamic thermal prediction and structural thermal environment prediction, and takes into account the influence of wall temperature changes on heat flux. Compared with the prediction method based on aerodynamic or structural multi-field coupling and the traditional structural thermal environment prediction technology using CFD aerodynamic thermal data, the present invention ensures higher accuracy, and the prediction time is almost equivalent to the aerodynamic thermal prediction of CFD, which improves the efficiency of the thermal structure design stage, completes the thermal environment prediction of the entire flight envelope, and greatly shortens and reduces the redundancy of the aircraft thermal structure design; the present invention uses the constructed heat flux correction model affected by the wall temperature to correct the heat flux value under isothermal wall conditions, and realizes the heat flux prediction correction under different wall temperature conditions at different positions under the aerodynamic heating effect of the wall. Without increasing the amount of CFD numerical calculation, the accuracy of the prediction data is improved. Compared with the traditional isothermal wall with a constant temperature, the corrected aerodynamic thermal data is closer to the actual flight condition. BRIEF DESCRIPTION OF THE DRAWINGS

[0064] The drawings described herein are used to provide a further understanding of the present invention and constitute a part of the present invention. The exemplary embodiments of the present invention and their descriptions are used to explain the present invention and do not constitute an improper limitation of the present invention. In the drawings:

[0065] Figure 1 A schematic diagram of a method for rapidly predicting the thermal environment of an aircraft skin structure along a flight envelope;

[0066] Figure 2 The present invention is a schematic flow chart of an embodiment of a method for rapidly predicting the thermal environment of an aircraft skin structure along a flight envelope;

[0067] Figure 3 It is a curve diagram of the modified model of heat flux affected by wall temperature;

[0068] Figure 4 It is a schematic diagram of the heat flux curve changing with the wall temperature in the laminar flow state of the 7-degree cone model;

[0069] Figure 5 It is a schematic diagram of the curve of heat flux changing with wall temperature in the turbulent state of the 7-degree cone model;

[0070] Figure 6 It is a schematic diagram of the adiabatic temperature variation curve of the wall of the 7-degree cone model;

[0071] Figure 7 It is a schematic diagram of the heat flux curve changing with the wall temperature at the x=0.1 position of the 7-degree cone model;

[0072] Figure 8 It is a schematic diagram of the heat flux curve changing with the wall temperature at the x=0.4 position of the 7-degree cone model;

[0073] Fig. 9 Schematic diagram of the heat flux curve varying with the wall temperature at the x=0.7 position for the 7-degree cone model. DETAILED DESCRIPTION

[0074] In order to make the technical solutions and advantages of the embodiments of the present invention more clearly understood, the exemplary embodiments of the present invention are further described in detail below in conjunction with the accompanying drawings. Obviously, the described embodiments are only part of the embodiments of the present invention, rather than an exhaustive list of all the embodiments. It should be noted that the embodiments of the present invention and the features in the embodiments can be combined with each other without conflict.

[0075] refer to Figure 1-Figure 9 The present embodiment is described in detail. A method for rapidly predicting the thermal environment of an aircraft skin structure along a flight envelope specifically comprises the following steps:

[0076] S1. Discretize the flight envelope to form a flight envelope point set for subsequent propulsion calculation along the time dimension of the flight envelope, that is, use discontinuity points to describe the dynamic flight process of the aircraft along the flight envelope;

[0077] Specifically: the flight state of each state point of the flight envelope point set includes flight time t, altitude H, Mach number Ma, angle of attack α, fluid pressure p, fluid density ρ and wall temperature T;

[0078] S2. Carry out aerodynamic thermal environment prediction and use CFD numerical calculation method to perform numerical simulation of aerodynamic thermal data under the flight state of the flight envelope point set;

[0079] Specifically: Aerothermal data include wall heat flux under isothermal wall boundary conditions and adiabatic wall temperature under adiabatic wall boundary conditions;

[0080] S3. constructing a correction model of heat flux affected by wall temperature by establishing a linear or quasi-linear temperature relationship according to heat flux values ​​under isothermal wall boundary conditions with different wall temperatures and adiabatic wall temperatures under adiabatic wall boundary conditions;

[0081] S4. Carry out thermal environment simulation of the aircraft skin structure along the flight envelope, correct the extracted aircraft skin structure temperature through the correction model of heat flux affected by wall temperature, and repeat the advancement calculation along the time dimension of the flight envelope until the thermal environment data of the skin structure at all state points under the flight envelope are obtained to form the thermal environment distribution of the skin structure along the flight envelope, so as to realize rapid prediction of the thermal environment.

[0082] Further, in said S2, aerodynamic thermal numerical simulation is performed by solving the three-dimensional unsteady compressible NS (Navier-Stokes) equations, and the body force and external heat source are not considered in the solution process, and the three-dimensional unsteady compressible NS equations are dimensionless;

[0083] The dimensionless form of the differential form of the three-dimensional unsteady compressible NS equations with respect to the conserved variable Q in the Cartesian coordinate system is expressed as:

[0084]

[0085] Where x is the horizontal axis of the Cartesian coordinate system, y is the vertical axis of the Cartesian coordinate system, z is the vertical axis of the Cartesian coordinate system, F is the inviscid vector flux of the first convection term, G is the inviscid vector flux of the second convection term, H is the inviscid vector flux of the third convection term, and F ν is the first viscous flux vector, G ν is the second viscous flux vector, H ν is the third viscous flux vector;

[0086] The conserved variable Q is expressed as:

[0087]

[0088] Wherein, u is the first velocity component in the Cartesian coordinate system, v is the second velocity component in the Cartesian coordinate system, w is the third velocity component in the Cartesian coordinate system, and E is the total energy per unit mass of gas;

[0089] The first convection term, the inviscid flux vector F, is expressed as:

[0090]

[0091] The second convection term, the inviscid flux vector G, is expressed as:

[0092]

[0093] The third convection term, the inviscid vector flux H, is expressed as:

[0094]

[0095] First viscous flux vector F ν It is expressed as:

[0096]

[0097]

[0098]

[0099]

[0100] Among them, τ xx is the first viscous stress term of the first viscous vector flux, τ xy is the second viscous stress term of the first viscous vector flux, τ xz is the third viscous stress term of the first viscous flux vector, is the fourth viscous stress term of the first viscous flux vector, q x is the heat flow in the x direction;

[0101] The second viscous flux G ν It is expressed as:

[0102]

[0103]

[0104]

[0105]

[0106] Among them, τ yx is the first viscous stress term of the second viscous vector flux, τ yy is the second viscous stress term of the second viscous vector flux, τ yz is the third viscous stress term of the second viscous flux vector, is the fourth viscous stress term of the second viscous vector flux, q y is the heat flow in the y direction;

[0107] The third viscous flux vector H ν It is expressed as:

[0108]

[0109]

[0110]

[0111]

[0112] Among them, τ zx is the first viscous stress term of the third viscous vector flux, τ zy is the second viscous stress term of the third viscous vector flux, τ zz is the third viscous stress term of the third viscous vector flux, is the fourth viscous stress term of the third viscous vector flux, q z is the heat flow in z direction;

[0113] The total energy E of a unit mass of gas is expressed as:

[0114]

[0115]

[0116] Among them, e is the internal energy per unit mass of gas, γ is the specific heat ratio;

[0117] Wall heat flux q w It is expressed as:

[0118]

[0119]

[0120] Where k is the convective heat transfer coefficient, μ is the molecular viscosity coefficient, c p is the specific heat capacity at constant pressure, r is the gas constant, is the normal unit length outside the wall, and the partial derivative of temperature and the normal unit length outside the wall Through T x , T y and T z Solve, T x , T y and T z represent the temperature gradients in the x, y, and z directions, respectively;

[0121] T x , T y and T z Respectively expressed as:

[0122] T x =T ξ ξ x +Tη η x +T ζ ζ x

[0123] T y =T ξ ξ y +T η η y +T ζ ζ y

[0124] T z =T ξ ξ z +T η η z +T ζ ζ z

[0125] Among them, ξ i , η i and i All represent coordinate conversion coefficients, i = x, y, z, T ξ , T η and T ζ They represent the temperature gradients in the directions of ξ, η and ζ respectively, which can be obtained from the solutions in the flow field and the temperature conditions at the boundaries;

[0126] Specifically, in the CFD numerical simulation process of solving the NS equations, all physical quantities were dimensionally processed. The dimensionless method was used to transform the control equations into a standard form, which can achieve similar simulation of the calculation. After dimensionless conversion, similar parameters such as the Reynolds number, Mach number and Prandtl number were obtained. The results of the above calculations are not only limited to a specific geometric shape, but also because of the use of characteristic scales, numerical errors caused by large or small actual scales are avoided.

[0127] Furthermore, in S3, the correction model of heat flux affected by wall temperature is expressed as:

[0128]

[0129] Among them, q w (T_w) is the heat flux value under isothermal wall boundary conditions when the wall temperature is T_w, T_ad is the adiabatic wall temperature, q w (T_P) is the heat flux value when the wall temperature is T_P.

[0130] Furthermore, in the S4, the heat flux value of the first state point of the flight envelope point set is passed as input to the heat conduction calculation program, and the advancement calculation along the time dimension of the flight envelope is carried out until the calculation is completed and the second state point is reached. According to the extracted skin structure surface temperature distribution data, the heat flux value of each point of the aircraft surface grid is corrected and calculated through a correction model of heat flux affected by wall temperature to obtain the corrected heat flux value of the second state point, and the corrected heat flux data is passed to the heat conduction calculation program, and the advancement calculation along the time dimension of the flight envelope is carried out until the calculation is completed and the next state point is reached. The above steps are repeated, and the solution is successively advanced until the calculation of all state points is completed, and the thermal environment distribution of the aircraft skin structure along the flight envelope is obtained;

[0131] Specifically, refer to Figure 2-9 , w / m 2 is the wall heat flux q w The unit is K, the unit of the wall temperature T, m is the unit of the specific position of the 7-degree cone model on the horizontal axis x of the Cartesian coordinate system, lami is the laminar flow calculation result, Turb is the turbulent flow calculation result, MODLE_lami is the laminar flow correction result, MODLE_Turb is the turbulent flow correction result, CFD_lami is the CFD laminar flow calculation result, CFD_Turb is the CFD turbulent flow calculation result. The present invention is applied to the design and development of hypersonic aircraft. For the prediction needs of the aerodynamic thermal environment and the structural thermal environment, a single machine can be used or a single machine combined with a supercomputing platform to achieve rapid prediction of the thermal environment of the aircraft skin structure along the flight envelope, and the calculation efficiency and accuracy are greatly improved under the condition of ensuring a certain calculation accuracy.

[0132] Although the present invention has been described according to a limited number of embodiments, it will be apparent to those skilled in the art, with the benefit of the above description, that other embodiments may be envisioned within the scope of the invention thus described. In addition, it should be noted that the language used in this specification is selected primarily for readability and teaching purposes, rather than for explaining or defining the subject matter of the present invention. Therefore, many modifications and variations will be apparent to those skilled in the art without departing from the scope and spirit of the appended claims. The disclosure of the present invention is illustrative, not restrictive, with respect to the scope of the present invention, which is defined by the appended claims.

Claims

1. A method for rapidly predicting the thermal environment of an aircraft skin structure along a flight envelope, characterized in that: The following steps are involved: S1. Discretize the flight envelope to form a flight envelope point set for subsequent propulsion calculation along the time dimension of the flight envelope, that is, use discontinuity points to describe the dynamic flight process of the aircraft along the flight envelope; Specifically: the flight state of each state point of the flight envelope point set includes flight time t, altitude H, Mach number Ma, angle of attack α, fluid pressure p, fluid density ρ and wall temperature T; S2. Carry out aerodynamic thermal environment prediction and use CFD numerical calculation method to perform numerical simulation of aerodynamic thermal data under the flight state of the flight envelope point set; Specifically: Aerothermal data include wall heat flux under isothermal wall boundary conditions and adiabatic wall temperature under adiabatic wall boundary conditions; S3. constructing a correction model of heat flux affected by wall temperature by establishing a linear or quasi-linear temperature relationship according to heat flux values ​​under isothermal wall boundary conditions with different wall temperatures and adiabatic wall temperatures under adiabatic wall boundary conditions; S4. Carry out thermal environment simulation of the aircraft skin structure along the flight envelope, correct the extracted aircraft skin structure temperature through the correction model of heat flux affected by wall temperature, and repeat the advancement calculation along the time dimension of the flight envelope until the thermal environment data of the skin structure at all state points under the flight envelope are obtained to realize rapid prediction of the thermal environment.

2. A method for rapid prediction of thermal environment of aircraft skin structure along the flight envelope according to claim 1, characterized in that: In S2, aerodynamic thermal numerical simulation is performed by solving the three-dimensional unsteady compressible NS equations, and the body force and external heat source are not considered in the solution process, and the three-dimensional unsteady compressible NS equations are dimensionless; The dimensionless form of the differential form of the three-dimensional unsteady compressible NS equations with respect to the conserved variable Q in the Cartesian coordinate system is expressed as: Where x is the horizontal axis of the Cartesian coordinate system, y is the vertical axis of the Cartesian coordinate system, z is the vertical axis of the Cartesian coordinate system, F is the inviscid vector flux of the first convection term, G is the inviscid vector flux of the second convection term, H is the inviscid vector flux of the third convection term, and F ν is the first viscous flux vector, G ν is the second viscous flux vector, H ν is the third viscous flux vector; The conserved variable Q is expressed as: Wherein, u is the first velocity component in the Cartesian coordinate system, v is the second velocity component in the Cartesian coordinate system, w is the third velocity component in the Cartesian coordinate system, and E is the total energy per unit mass of gas; The first convection term, the inviscid flux vector F, is expressed as: The second convection term, the inviscid flux vector G, is expressed as: The third convection term, the inviscid vector flux H, is expressed as: First viscous flux vector F ν It is expressed as: Among them, τ xx is the first viscous stress term of the first viscous vector flux, τ xy is the second viscous stress term of the first viscous vector flux, τ xz is the third viscous stress term of the first viscous flux vector, is the fourth viscous stress term of the first viscous flux vector, q x is the heat flow in the x direction; The second viscous flux vector G ν It is expressed as: Among them, τ yx is the first viscous stress term of the second viscous vector flux, τ yy is the second viscous stress term of the second viscous vector flux, τ yz is the third viscous stress term of the second viscous flux vector, is the fourth viscous stress term of the second viscous flux vector, q y is the heat flow in the y direction; The third viscous flux vector H ν It is expressed as: Among them, τ zx is the first viscous stress term of the third viscous vector flux, τ zy is the second viscous stress term of the third viscous vector flux, τ zz is the third viscous stress term of the third viscous vector flux, is the fourth viscous stress term of the third viscous vector flux, q z is the heat flow in z direction; The total energy E of a unit mass of gas is expressed as: Among them, e is the internal energy per unit mass of gas, γ is the specific heat ratio; Wall heat flux q w It is expressed as: Where k is the convective heat transfer coefficient, μ is the molecular viscosity coefficient, c p is the specific heat capacity at constant pressure, r is the gas constant, is the normal unit length outside the wall, and the partial derivative of temperature and the normal unit length outside the wall By T x 、T y and T z Solve, T x 、T y and T z represent the temperature gradients in the x, y, and z directions, respectively; T x , T y and T z Respectively expressed as: T x =T ξ x x +T η or x +T ζ g x T y =T ξ x y +T η or y +T ζ g y T z =T ξ x z +T η or z +T ζ g z Among them, ξ i , η i and i All represent coordinate conversion coefficients, i = x, y, z, T ξ , T η and T ζ They represent the temperature gradients in the directions of ξ, η and ζ respectively.

3. The method for rapid prediction of thermal environment of aircraft skin structure along the flight envelope according to claim 2, characterized in that: In S3, the heat flux is expressed by the modified model of wall temperature as follows: Among them, q w (T_w) is the heat flux value under isothermal wall boundary conditions when the wall temperature is T_w, T_ad is the adiabatic wall temperature, q w (T_P) is the heat flux value when the wall temperature is T_P.

4. A method for rapid prediction of thermal environment of aircraft skin structure along the flight envelope according to claim 3, characterized in that: In the S4, the heat flux value of the first state point of the flight envelope point set is passed as input to the heat conduction calculation program, and the advancement calculation is carried out along the time dimension of the flight envelope until the calculation is completed and the second state point is reached. According to the extracted skin structure surface temperature distribution data, the heat flux value of each point of the aircraft surface grid is corrected and calculated through a correction model of heat flux affected by wall temperature to obtain the corrected heat flux value of the second state point, and the corrected heat flux data is passed to the heat conduction calculation program, and the advancement calculation is carried out along the time dimension of the flight envelope until the calculation is completed and the next state point is reached. The above steps are repeated, and the solution is successively advanced until the calculation of all state points is completed, and the thermal environment distribution of the aircraft skin structure along the flight envelope is obtained.

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