Method for simulating transient heat and mass transfer processes on aircraft surfaces based on zonal modeling and interface coupling

Abstract

The application discloses a method for simulating transient heat and mass transfer process of an aircraft surface based on partition modeling and interface coupling, separates and solves high-speed compressible flow of an outer flow area of the aircraft and low-speed flow of an inner flow area of the aircraft based on the idea of 'partition modeling and interface coupling', and designs a transmission method between data of the inner flow area and the outer flow area, so that transient coupling solution of the inner flow field and the outer flow field of the aircraft is realized. The application comprehensively considers the flow of the coolant in the multi-hole shell of the aircraft under the operating condition and the non-steady heat and mass transfer process between the coolant and the high-Mach-number flow field outside the multi-hole shell, can accurately predict temperature and heat flow distribution of the surface structure of the aircraft under the operating condition, and has guiding significance for fine design of the surface thermal structure and the cooling system of the aircraft.

Description

Technical Field

[0001] This invention belongs to the field of aircraft thermal management technology, specifically relating to a method for simulating transient heat and mass transfer processes on the surface of an aircraft based on partition modeling and interface coupling. Background Technology

[0002] High-speed flight technology is considered a disruptive technology of the future, which will revolutionize the design concept of manned spacecraft. Future high-speed aircraft will approach the limits of long-duration maneuvering flight within the atmosphere, possessing capabilities such as long operating time and high speed. However, such aircraft experience extremely severe aerodynamic heating during prolonged, high-speed flight in the thin atmosphere, necessitating the application of cooling technologies to eliminate the impact of aerodynamic heat on internal electronic components. Therefore, developing compact, high-efficiency cooling technologies has become a key aspect of the design and development of such aircraft. Among existing cooling technologies, sweating cooling technology possesses high cooling capacity and requires relatively small amounts of coolant, attracting significant attention from academia and industry in recent years. Optimizing the sweating cooling system to precisely match the distribution characteristics of the aircraft's surface heat load has become crucial for improving the aircraft's thermal protection capabilities. However, before optimizing the design, it is essential to obtain the transient changes in surface temperature and heat flow under real flight conditions to identify the root causes of problems in the existing system. Considering the complexity of flight experiments, numerical simulation is a more feasible method to simulate the heat and mass transfer process in the internal and external flow regions of an aircraft, thereby obtaining the transient temperature and heat flux distribution changes on the aircraft surface. However, a series of challenges remain in the numerical simulation of heat and mass transfer processes in the internal and external flow regions of an aircraft. These include: (1) the lack of a calculation method that couples the high Mach number flow outside the aircraft with the low-speed flow inside; and (2) the aircraft's speed and altitude will change over time during operation, but there is currently a lack of methods to study the impact of transient changes in the external flow field on the internal flow field.

[0003] Currently, scholars both domestically and internationally have conducted a series of studies on the heat and mass transfer process on aircraft surfaces considering the influence of external flow fields. Dong and Wang used a semi-mixed phase model to conduct numerical studies on the heat and mass transfer process in a sweating cooling system and analyzed the influence of porous framework thermal conductivity and porosity on the system's cooling capacity. This paper applies non-uniform heat flux boundary conditions at the heated end of the internal flow region to reflect the influence of the external flow field on the flow heat transfer process in the internal flow region. Lin Boxi et al. developed a multi-physical coupling system for simulating steady-state sweating cooling, which can be used to analyze the heat and mass transfer process inside and outside the aircraft structure under steady-state service conditions. Su et al. combined a two-phase mixing model to study the heat and mass transfer process in a sweating cooling system of an aircraft under steady-state conditions and analyzed the influence of parameters such as porous structure particle diameter on cooling efficiency. From the above analysis, it can be seen that existing studies on the heat and mass transfer process on aircraft surfaces often simplify the simulation of unsteady high Mach number external flow fields to varying degrees, making it difficult to study the influence of transiently changing external flow fields on the internal flow field. Summary of the Invention

[0004] To address the shortcomings of current numerical simulation methods for heat and mass transfer processes on aircraft surfaces, this invention proposes a method based on partitioned modeling and interface coupling to simulate transient heat and mass transfer processes on aircraft surfaces. This method can obtain the transient changes in temperature and heat flux distribution on the aircraft surface within a time-varying external flow field, thus aiding in the refined design of aircraft surface cooling systems.

[0005] The method proposed in this invention includes the following steps:

[0006] A method for simulating transient heat and mass transfer processes on the surface of an aircraft based on partitioned modeling and interface coupling, applicable to aircraft employing a phase change sweating cooling system, includes the following steps:

[0007] Step 1: Based on the actual structure of the aircraft, physical models of the internal flow region on the surface of the aircraft and the internal and external flow regions in the adjacent space are established respectively. Then, the internal and external flow regions are divided into meshes to achieve the discretization of the physical space. Based on the convergence of heat flow and pressure at the interface between the internal and external flow regions, the mesh division of the internal and external flow regions is checked for independence.

[0008] Step 2: Based on the actual operating conditions of the aircraft, determine the static temperature and static pressure of the outflowing air as the initial conditions of the outflowing region, and determine the initial temperature, pressure and flow distribution of the interface between the inflowing and outflowing regions and the inflowing region based on the temperature, pressure and flow rate of the fluid in the inflowing region.

[0009] Step 3: Taking into account the flow, diffusion, and phase change processes of the coolant overflowing from the inner flow region in the outer flow region, and the compressibility of the air in the outer flow region, the k-ε model is used to describe the turbulence in the outer flow region. The multi-component fluid flow and heat transfer calculation model in the outer flow region is established as follows:

[0010]

[0011]

[0012]

[0013]

[0014]

[0015]

[0016]

[0017] In the formula, ρ is the density of the fluid in the outer flow region, and u is the velocity of the fluid in the outer flow region. i u j u k and u l Let i, j, k, and l represent the components of the fluid velocity in the outflow region along the directions i, j, k, and l, respectively, and x be the coordinates of the fluid in the outflow region. i x j x k and x l δ represents the components of the fluid coordinates in the i, j, k, and l directions, respectively; p is the pressure of the fluid in the outer flow region; T is the temperature of the fluid in the outer flow region; h is the specific enthalpy of the fluid in the outer flow region; E is the total energy of the fluid in the outer flow region; μ is the dynamic viscosity coefficient of the fluid in the outer flow region; and t is the flow time. ij It is a component of the unit second-order tensor, and its value is 1 if and only if i = j, otherwise it is 0; It is the Reynolds stress, where u′ i and u′ j Let be the fluctuating velocities of the fluid in the i and j directions within the outflow region, respectively; k, ε, and μ are the pulsating velocities of the fluid in the i and j directions, respectively. t These represent the turbulent kinetic energy, dissipation rate, and turbulent viscosity coefficient of the fluid, respectively; σ k and σ ε These are the Prandtl numbers for turbulent kinetic energy and dissipation rate, respectively; C1 and C2 are both constants; Y n D is the mass fraction of the nth component in the fluid in the outflow region. m,n and D T,n These are the mass diffusion coefficient and Soret diffusion coefficient of the nth component in the outflow region, respectively; h nand τ represents the specific enthalpy and diffusion flux of the nth component in the outflow region fluid, respectively; ij ) eff It is the deviatoric stress tensor of the fluid in the outflow region; S m and S h Representing the mass source term and energy source term in the coolant phase change process, respectively, the Lee model can be used to describe the above source terms as follows:

[0018]

[0019] S h =S m ·h0

[0020] In the formula, C evap T is the evaporation constant of the Lee model. sat h0 and k represent the phase change temperature and phase change enthalpy of the coolant, respectively; eff This represents the equivalent thermal conductivity of the fluid in the outflow region; the various physical properties of the fluid in the outflow region can be calculated using the following methods:

[0021]

[0022]

[0023]

[0024]

[0025] In the formula, R is the molar gas constant; c p M represents the specific isobaric heat capacity of the fluid in the outflow region; n c p,n k eff,n and μ n These are the relative molar mass, specific heat capacity at constant pressure, equivalent thermal conductivity, and dynamic viscosity coefficient of the nth component in the outflow region fluid, respectively.

[0026] The solution process of the above model is as follows: First, initial conditions are set. When solving the flow and heat transfer process in the outer flow region of the first time step, the distribution of physical quantities set in step 2 is used as the initial conditions; when not solving the first time step, the distribution of each physical quantity in the outer flow region of the previous time step is used as the initial conditions for solving this time step. Second, boundary conditions are set. The flow rate and temperature of the coolant overflowing from the inner flow region in the calculation results of the previous iteration step are used as the boundary conditions of the interface between the inner and outer flow regions; the Mach number of the outer flow and the static pressure and static temperature of the outer flow air are calculated based on the current speed and altitude of the spacecraft and used as the boundary conditions for the inlet and outlet of the outer flow region. Third, the finite volume method is used. Based on the laws of conservation of mass, momentum, and energy, the partial differential equations in the multi-component fluid flow and heat transfer calculation model in the outer flow region are converted into a linear difference equation system and solved. Finally, after the calculation converges, the temperature and pressure distribution of the outer flow region in this time step and this iteration step are obtained. Based on the calculation results, the pressure and heat flux density on the outer side of the interface between the internal and external flow regions are output. The method for calculating the heat flux density is as follows:

[0027]

[0028] In the formula, q o T represents the heat flux density on the outer side of the interface between the inflow and outflow regions. 1,o and T 0,o Here, represents the temperature at the center of the boundary grid and the center of the boundary surface, respectively, and ds represents the distance from the center of the boundary grid to the center of the boundary surface. and These are the volume vector of the boundary mesh and the unit normal vector of the boundary surface, respectively.

[0029] Step 4: Based on the calculation results of the external flow region, design an interaction method for the data at the interface between the internal and external flow regions; set the pressure and heat flux density of the fluid outside the interface between the internal and external flow regions as the boundary conditions for the flow heat transfer calculation in the internal flow region. The internal flow region on the aircraft surface is composed of a porous medium. The thermal diffusivity of the solid phase structure in the porous medium is much higher than that of the coolant. Therefore, under the condition of external heat flow, there is a huge difference in temperature and heat flux between the solid and liquid phases at the interface between the internal and external flow regions. Based on the ratio of the effective thermal conductivity of the solid and liquid phases, the heat flux density at the boundary between the solid and liquid phases of the porous medium in the internal flow region can be expressed as:

[0030]

[0031] In the formula, e represents the porosity of the porous medium in the internal region, and k s and k f q represents the thermal conductivity of the solid and liquid phases in a porous medium, respectively. i,f and q i,s q represents the heat flux density at the solid-phase boundary and liquid-phase boundary of the porous medium in the internal flow region, respectively, while q iThis represents the total heat flux density at the boundary of the inflow region.

[0032] Step 5: Taking into account the phase change process of the coolant in the internal flow region of the aircraft surface and the influence of the capillary structure of the porous medium on the liquid phase flow, a flow and heat transfer calculation model inside the aircraft surface structure is established by combining a local thermal non-equilibrium (LTNE) model with a two-phase mixture model (TPMM) as follows:

[0033]

[0034]

[0035]

[0036]

[0037] In the formula, γ is the velocity of the fluid within the porous medium in the inflow region, K is the permeability of the porous medium in the inflow region, μ is the dynamic viscosity coefficient of the coolant under two-phase mixing conditions, and γ is the velocity of the fluid within the porous medium in the inflow region. H and Γ H ρ represents the coefficients of the convection and diffusion terms in the enthalpy equation, respectively. s ,c p,s ,k s and T s These are the density, specific isobaric heat capacity, thermal conductivity, and temperature of the porous medium solid phase in the inflow region, respectively. sf It refers to the heat exchange between the solid and liquid phases of the porous medium in the inflow region. Let H represent divergence, and H be the enthalpy of the coolant in the internal flow region, expressed as follows:

[0038] H = ρh - H0, H0 = ρh v,sat

[0039] In the formula, H0 is the coolant enthalpy reference point set in the calculation, h is the specific enthalpy of the coolant, and h v,sat This is the saturated vapor enthalpy of the coolant.

[0040] The solution process of the above model is as follows: First, initial conditions are set. When solving the flow and heat transfer process in the inner flow region of the first time step, the distribution of physical quantities in the inner flow region set in step 2 is used as the initial conditions; when not solving the first time step, the distribution of physical quantities in the inner flow region of the previous time step is used as the initial conditions for solving this time step. Second, boundary conditions are set. The pressure and heat flux density of the fluid outside the interface between the inner and outer flow fields in the calculation results of the previous iteration step are used as the boundary conditions at the interface between the inner and outer flow regions. The temperature and set flow rate of the coolant inside the aircraft are used as the boundary conditions for the coolant inlet. Third, the partial differential equations in the flow and heat transfer calculation model inside the surface structure of the aircraft are converted into a linear difference equation system using the finite volume method and solved. Finally, after the calculation converges, the temperature and velocity distribution of the inner flow region in this time step and this iteration step are obtained. Based on the calculation results, the composition, temperature, and flow rate of the coolant inside the interface between the inner and outer flow fields are output, where the calculation methods for the coolant temperature and flow rate are as follows:

[0041] T i =eT f +(1-e)T s

[0042]

[0043] In the formula, T i It is the temperature on the inner boundary grid of the interface between the inflow and outflow regions, T f and T s q represents the temperature at the center point of the liquid and solid phase boundary grids, respectively. m,i It is the flow rate of coolant overflowing from the internal flow area.

[0044] Step 6: Repeat steps 3, 4, and 5. During the iterative process, the data at the interface between the internal and external flow regions continuously interact, and the difference between the physical quantities on both sides of the interface gradually decreases. When the five conditions of continuous temperature, continuous heat flux density, continuous pressure, continuous flow rate, and continuous composition are achieved on both sides of the interface between the internal and external flow regions, the iteration process terminates. The distribution of physical quantities in the internal and external flow regions in the last iteration is the distribution of physical quantities on the aircraft surface considering the coupling of the internal and external flow fields at the current time step.

[0045] Step 7: Repeat steps 3 to 6 to obtain the distribution of physical quantities on the surface of the aircraft at different time steps, and finally clarify the transient change process of physical quantities such as surface temperature and heat flux of the aircraft considering the coupling of internal and external flow fields during the entire service process.

[0046] Compared with the prior art, the beneficial effects of the present invention are:

[0047] 1) Based on the idea of ​​"partition modeling and interface coupling", the high Mach number flow of the external air and the low speed flow of the internal coolant of the aircraft are modeled and solved separately. According to the physical characteristics of the two regions, a more suitable solver and a discretization method for pressure-velocity coupling are assigned. An interface coupling algorithm is introduced to realize the continuity of temperature, pressure, flow rate, composition and heat flux density at the boundary of the two regions, and to clarify the transient change process of surface temperature and heat flux distribution of the aircraft under the influence of the external flow field.

[0048] 2) The impact of the time-varying nature of the external flow field of the aircraft on the heat and mass transfer process in the internal flow region during service was clarified; the calculation results can guide the refined design of the aircraft surface cooling system, realize the real-time and accurate matching between the coolant flow rate and the external transient heat load, so as to achieve the goal of high efficiency and lightweight cooling system. Attached Figure Description

[0049] Figure 1 This invention is a numerical method for simulating the transient heat and mass transfer process on the surface of an aircraft based on "partition modeling and interface coupling".

[0050] Figure 2 A physical model of the surface sweating cooling components of an aircraft under supersonic conditions with an angle of attack of 20° and a Mach number of 4.5.

[0051] Figure 3 Mesh generation of external flow field and surface sweating cooling components of aircraft

[0052] Figure 4 Verification of mesh independence of external flow field and surface sweating cooling components of aircraft

[0053] Figure 5 Transient variation of coolant flow distribution on the surface of an aircraft under supersonic conditions with an angle of attack of 20° and a Mach number of 4.5.

[0054] Figure 6 Transient change process of heat flux density distribution on the surface of the aircraft under supersonic mainstream conditions with an angle of attack of 20° and a Mach number of 4.5.

[0055] Figure 7 Transient change process of surface temperature distribution of aircraft under supersonic mainstream conditions with an angle of attack of 20° and a Mach number of 4.5.

[0056] Figure 8 Transient change process of pressure distribution on the surface of the aircraft under supersonic mainstream conditions with an angle of attack of 20° and a Mach number of 4.5. Detailed Implementation

[0057] The invention will now be described in detail below with reference to the accompanying drawings, taking a two-dimensional sweating cooling system in a supersonic flow field as an example:

[0058] like Figure 1 As shown, the method of simulating the transient heat and mass transfer process on the surface of an aircraft based on partition modeling and interface coupling is as follows:

[0059] First, the flat plate structure on the surface of the aircraft and its onboard phase change sweating cooling system, such as Figure 2 As shown. The aircraft surface structure consists of metal structural components and a porous medium region, with a water tank placed beneath the porous medium region. Under operational conditions, the supersonic current outside the aircraft passes over the flat plate surface, forming a shock wave. The airflow temperature behind the shock wave rises sharply. At this time, the coolant in the water tank undergoes a gas-liquid phase change and flows through the porous medium region to the outer flow region, achieving sweating cooling of the aircraft surface. Due to the morphological characteristics of the flat plate structure on the aircraft surface, this three-dimensional structure can be equivalent to a two-dimensional planar model to reduce computational resource requirements. Physical models of the inner flow region and the supersonic outer flow region of the flat plate sweating cooling system are established separately. Structured meshes are generated for both the inner and outer flow regions, and the meshes at the near-interface portions of the inner and outer flow regions are refined. The mesh generation results are shown below. Figure 3 As shown. Furthermore, based on the convergence of heat flux and pressure at the interface between the internal and external flow fields, the meshes of the internal and external flow regions are checked for independence, and the results are as follows. Figure 4 As shown. From Figure 4 As can be seen from the data, when the number of grids is higher than 235,800, the effect of further refining the grid on the calculation results of heat flux density and fluid pressure at the interface between the internal and external flow regions is less than 1%. Therefore, this set of grids was selected to simulate the transient heat and mass transfer process on the surface of the aircraft's flat plate structure.

[0060] Second, the static temperature and static pressure of the outflowing air were selected to be 226.5 K and 1172 Pa, respectively; water was selected as the coolant in the sweating cooling system, with an initial temperature of 300 K; and the coolant flow rate at the system inlet was set to 0.3 kg·m³. -2 ·s -1 Based on the above operating conditions, set the initial conditions for the internal and external flow areas;

[0061] Third, taking into account the flow, diffusion, and phase change processes of the coolant overflowing from the inner flow region in the outer flow region, and the compressibility of the air in the outer flow region, the k-ε model is used to describe the turbulence of the fluid in the outer flow region, and the following calculation model for the multi-component fluid flow and heat transfer in the outer flow region is established:

[0062]

[0063]

[0064]

[0065]

[0066]

[0067]

[0068]

[0069] In the formula, ρ is the density of the fluid in the outer flow region, and u is the velocity of the fluid in the outer flow region. i u j u k and u l Let i, j, k, and l represent the components of the fluid velocity in the outflow region along the directions i, j, k, and l, respectively, and x be the coordinates of the fluid in the outflow region. i x j x k and x l δ represents the components of the fluid coordinates in the i, j, k, and l directions, respectively; p is the pressure of the fluid in the outer flow region; T is the temperature of the fluid in the outer flow region; h is the specific enthalpy of the fluid in the outer flow region; E is the total energy of the fluid in the outer flow region; μ is the dynamic viscosity coefficient of the fluid in the outer flow region; and t is the flow time. ij It is a component of the unit second-order tensor, and its value is 1 if and only if i = j, otherwise it is 0; It is the Reynolds stress, where u′ i and u′ j Let be the fluctuating velocities of the fluid in the i and j directions within the outflow region, respectively; k, ε, and μ are the pulsating velocities of the fluid in the i and j directions, respectively. t These represent the turbulent kinetic energy, dissipation rate, and turbulent viscosity coefficient of the fluid, respectively; σ k and σ ε These are the Prandtl numbers for turbulent kinetic energy and dissipation rate, respectively; C1 and C2 are both constants; Y n D is the mass fraction of the nth component in the fluid in the outflow region. m,n and D T,n These are the mass diffusion coefficient and Soret diffusion coefficient of the nth component in the outflow region, respectively; h n and τ represents the specific enthalpy and diffusion flux of the nth component in the outflow region fluid, respectively; ij ) eff It is the deviatoric stress tensor of the fluid in the outflow region; S m and S h Representing the mass source term and energy source term in the coolant phase change process, respectively, the Lee model can be used to describe the above source terms as follows:

[0070]

[0071] S h =S m ·h0

[0072] In the formula, C evap T is the evaporation constant of the Lee model. sat h0 and k represent the phase change temperature and phase change enthalpy of the coolant, respectively; eff This represents the equivalent thermal conductivity of the fluid in the outflow region; the various physical properties of the fluid in the outflow region can be calculated using the following methods:

[0073]

[0074]

[0075]

[0076]

[0077] In the formula, R is the molar gas constant; c p M represents the specific isobaric heat capacity of the fluid in the outflow region; n c p,n k eff,n and μ n These are the relative molar mass, specific heat capacity at constant pressure, equivalent thermal conductivity, and dynamic viscosity coefficient of the nth component in the outflow region fluid, respectively.

[0078] The solution process of the above model is as follows: First, initial conditions are set. When solving the flow and heat transfer process in the outer flow region of the first time step, the physical quantity distribution set in step 2 is used as the initial conditions; when not solving the first time step, the distribution of each physical quantity in the outer flow region of the previous time step is used as the initial conditions for solving this time step. Second, boundary conditions are set. The flow rate and temperature of the coolant overflowing from the inner flow region in the calculation results of the previous iteration step are used as the boundary conditions of the interface between the inner and outer flow regions; the Mach number of the outer flow and the static pressure and static temperature of the outer flow air are calculated based on the current speed and altitude of the spacecraft and used as the boundary conditions for the inlet and outlet of the outer flow region. Third, the finite volume method is used. Based on the laws of conservation of mass, momentum, and energy, the partial differential equations in the multi-component fluid flow and heat transfer calculation model in the outer flow region are converted into a linear difference equation system and solved. Finally, after the calculation converges, the temperature and pressure distribution of the outer flow region in this time step and this iteration step are obtained. Based on the calculation results, the pressure and heat flux density on the outer side of the interface between the internal and external flow regions are output. The method for calculating the heat flux density is as follows:

[0079]

[0080] In the formula, q o T represents the heat flux density on the outer side of the interface between the inflow and outflow regions. 1,o and T 0,oHere, represents the temperature at the center of the boundary grid and the center of the boundary surface, respectively, and ds represents the distance from the center of the boundary grid to the center of the boundary surface. and These are the volume vector of the boundary mesh and the unit normal vector of the boundary surface, respectively.

[0081] The discretization and solution of the partial differential equations in the multi-component fluid flow and heat transfer calculation model in the above-mentioned outflow region can be completed using the density-based solver in ANSYS FLUENT software. During the solution process, the physical property settings of the fluid in the outflow region, the loading of flow boundary conditions at the interface between the inflow and outflow regions, and the output and saving of the calculation results are achieved through user-defined functions (UDFs) in FLUENT.

[0082] Fourth, based on the pressure and heat flux density of the fluid at the interface between the inner and outer flow regions in the calculation results of the outer flow region, the boundary conditions for the flow heat transfer calculation in the inner flow region are obtained. The inner flow region on the aircraft surface is composed of a porous medium. The thermal diffusivity of the solid phase structure in the porous medium is much higher than that of the coolant. Therefore, under the condition of external heat flow, there is a huge difference in temperature and heat flux between the solid and liquid phases at the interface between the inner and outer flow regions. Based on the ratio of the effective thermal conductivity of the solid and liquid phases, the heat flux density at the boundary between the solid and liquid phases of the porous medium in the inner flow region can be expressed as:

[0083]

[0084] In the formula, e represents the porosity of the porous medium in the internal region, and k s and k f q represents the thermal conductivity of the solid and liquid phases in a porous medium, respectively. i,f and q i,s q represents the heat flux density at the solid-phase boundary and liquid-phase boundary of the porous medium in the internal flow region, respectively, while q i This represents the total heat flux density at the boundary of the inflow region.

[0085] In this step, the data interaction method at the interface between the internal and external flow regions can be implemented based on the AAS (Asa Server) module in Matlab software. Through this module, the interaction of temperature, pressure, flow rate, composition and heat flux density on both sides of the interface between the internal and external flow regions can be realized, and the continuity of physical quantities on both sides of the interface between the internal and external flow regions can be determined, thereby helping to simulate the transient heat and mass transfer process on the surface of the aircraft.

[0086] Fifth, taking into account the coolant seepage and phase change processes in the internal flow region of the aircraft surface, as well as the influence of the capillary structure of the porous medium on the liquid phase flow, a non-thermal equilibrium model and a two-phase mixing model are adopted to establish the flow and heat transfer calculation model inside the aircraft surface structure as follows:

[0087]

[0088]

[0089]

[0090]

[0091] In the formula, γ is the velocity of the fluid within the porous medium in the inflow region, K is the permeability of the porous medium in the inflow region, μ is the dynamic viscosity coefficient of the coolant under two-phase mixing conditions, and γ is the velocity of the fluid within the porous medium in the inflow region. H and Γ H ρ represents the coefficients of the convection and diffusion terms in the enthalpy equation, respectively. s ,c p,s ,k s and T s These are the density, specific isobaric heat capacity, thermal conductivity, and temperature of the porous medium solid phase in the inflow region, respectively. sf It refers to the heat exchange between the solid and liquid phases of the porous medium in the inflow region. Let H represent divergence, and H be the enthalpy of the coolant in the internal flow region, expressed as follows:

[0092] H = ρh - H0, H0 = ρh v,sat

[0093] In the formula, H0 is the coolant enthalpy reference point set in the calculation, h is the specific enthalpy of the coolant, and h v,sat This is the saturated vapor enthalpy of the coolant.

[0094] The solution process of the above model is as follows: First, initial conditions are set. When solving the flow and heat transfer process in the inner flow region of the first time step, the physical quantity distribution set in step 2 is used as the initial conditions; when not solving the first time step, the distribution of each physical quantity in the inner flow region of the previous time step is used as the initial conditions for solving this time step. Second, boundary conditions are set, that is, the pressure and heat flux density of the fluid outside the interface between the inner and outer flow fields in the calculation results of the previous iteration step are used as boundary conditions. Third, the partial differential equations in the flow and heat transfer calculation model inside the surface structure of the aircraft are converted into a linear difference equation system using the finite volume method, and then solved. Finally, after the calculation converges, the temperature and velocity distribution of the inner flow region in this time step and this iteration step are obtained. Based on the calculation results, the composition, temperature, and flow rate of the coolant inside the interface between the inner and outer flow fields are output, where the calculation methods for the coolant temperature and flow rate are as follows:

[0095] T i =eT f +(1-e)T s

[0096]

[0097] In the formula, T i It is the temperature on the inner boundary grid of the interface between the inflow and outflow regions, T f and T s q represents the temperature at the center point of the liquid and solid phase boundary grids, respectively. m,i It is the flow rate of coolant overflowing from the internal flow area.

[0098] Similarly, the discretization and solution of the partial differential equations in the flow heat transfer calculation model inside the surface structure of the aircraft can be completed with the pressure-based solver in ANSYS FLUENT software. During the solution process, the setting of coolant properties, the loading of heat flux density boundary conditions at the interface between the internal and external flow fields, and the output and saving of calculation results can all be achieved through user-defined functions in FLUENT.

[0099] Sixth, during the continuous iteration process, the difference between the physical quantities on both sides of the interface between the internal and external flow regions gradually narrows. When the five conditions of continuous temperature, continuous heat flux density, continuous pressure, continuous flow rate, and continuous composition are achieved on both sides of the interface between the internal and external flow regions, the iteration process terminates. The distribution of physical quantities in the internal and external flow regions in the last iteration is the distribution of physical quantities on the aircraft surface considering the coupling of the internal and external flow fields at the current time step.

[0100] Seventh, calculate the transient changes in the distribution of physical quantities on the surface of the flat plate sweating cooling system during the 12s time interval of the coupling of internal and external flow fields. The distributions of coolant flow rate, heat flux, temperature, and pressure on the system surface are as follows: Figure 5 , Figure 6 , Figure 7 and Figure 8 As shown, the transient changes in the surface temperature distribution of the aircraft indicate that within 12 seconds, the maximum surface temperature does not exceed 378 K, and after 6 seconds of contact between the internal and external fields, the increase in the maximum surface temperature is less than 1.1% with the continued extension of the operating time. These results demonstrate that the water-based evaporative cooling system can meet the cooling requirements of the aircraft's flat plate structure at 4.5 Mach. However, the transient distribution of fluid pressure, flow rate, and heat flux density on the porous medium surface shows that after 6 seconds of contact between the internal and external fields, the coolant on the near-sonic inflow side of the porous medium has completely vaporized, with the local coolant flow rate dropping to as low as 0.024 kg·m³. -2 ·s -1This leads to a decline in the overall performance of the sweating cooling system. The cause of this heat transfer deterioration is likely the sudden pressure increase in the porous medium near the incoming flow area due to coolant vaporization, which obstructs coolant flow. Consequently, the coolant flows to areas within the porous medium farther from the incoming flow, with lower heat flux density and lower pressure, resulting in uneven coolant distribution within the sweating cooling system. To further improve the efficiency of the sweating cooling system for the aircraft's flat plate structure, a matching design for the transient distribution of coolant flow should be proposed, based on the simulation results and the transient distribution of heat flux on the porous medium surface. Specifically, based on the simulated heat flux distribution, a coolant supply strategy that varies with space-time parameters should be designed. When heat transfer deterioration occurs in the sweating cooling system (9 seconds after operation), the coolant flow rate in areas with higher heat flux density should be increased to maintain a uniform coolant flow distribution within the porous medium under aerodynamic heating conditions, thus preemptively eliminating the impact of heat transfer deterioration within the porous medium on the overall performance of the sweating cooling system. Unlike traditional sweating cooling systems that rely on spatial variations in coolant supply, this patented coolant supply strategy, which adapts to changes in space and time parameters, can match the unsteady heat flux of the aircraft. When the heat flux density is low (e.g., within 9 seconds of operation), the coolant supply can be appropriately reduced, while still meeting the cooling requirements of the aircraft's surface structure.

[0101] As can be seen, this invention proposes a numerical method based on "partitioned modeling and interface coupling" to simulate the transient heat and mass transfer process on the aircraft surface. This method can accurately obtain the transient temperature and heat flux density changes inside the sweating cooling system of an aircraft surface during actual service, and clarify the spatiotemporal non-uniform characteristics of the heat load distribution on the aircraft surface under aerodynamic heating conditions. Based on this, a coolant flow distribution strategy that matches the surface heat load distribution characteristics can be rationally designed, improving the coolant utilization efficiency within the sweating cooling system, reducing the total coolant requirement of the system, and achieving the lightweight design goal of the sweating cooling system. This invention can accurately predict the time and location of heat transfer deterioration phenomena inside the sweating cooling system, providing guidance for the efficient and lightweight design of aircraft surface cooling systems.

[0102] This invention, based on the concept of "partitioned modeling and interface coupling," separately solves the high Mach number flow in the outer flow region and the low-speed flow in the inner flow region of an aircraft. It also designs a method for transferring data from the interface between the inner and outer flow regions, thereby achieving transient coupled solution of the aircraft's internal and external flow fields. This invention can accurately predict the transient temperature and heat flux distribution of the aircraft's surface structure under real outer flow conditions during operation, providing guidance for the refined design of the aircraft's surface thermal structure and cooling system.

Claims

1. A method for simulating transient heat and mass transfer processes on the surface of an aircraft based on partitioned modeling and interface coupling, applicable to aircraft employing a phase change sweating cooling system, wherein the internal flow region of the aircraft surface is composed of a porous medium, and the thermal diffusivity of the solid phase structure in the porous medium is much higher than that of the coolant, characterized in that... Includes the following steps: Step 1) Based on the actual structure of the aircraft, establish physical models of the internal flow region on the surface of the aircraft and the internal and external flow regions in the adjacent space. Then, divide the internal and external flow regions into grids to achieve the discretization of the physical space. Based on the convergence of heat flow and pressure at the interface between the internal and external flow regions, check the independence of the grid division of the internal and external flow regions. Step 2) Based on the actual operating conditions of the aircraft, determine the static temperature and static pressure of the outflowing air as the initial conditions of the outflowing region, and determine the initial temperature, pressure and flow distribution of the interface between the inflowing and outflowing regions and the inflowing region based on the temperature, pressure and flow rate of the fluid in the inflowing region. Step 3) Taking into account the flow, diffusion and phase change process of the coolant overflowing from the inner flow region in the outer flow region, as well as the compressibility of the air in the outer flow region, the k-ε model is used to describe the turbulence in the outer flow region. A multi-component fluid flow and heat transfer calculation model is established in the outer flow region to solve the transient physical quantity distribution in the outer flow region of the aircraft. Step 4) Based on the calculation results of the outflow area, design a data interaction method at the interface between the inflow and outflow areas; The pressure and heat flux density of the fluid outside the interface between the inner and outer flow regions are set as the boundary conditions for the flow heat transfer calculation in the inner flow region. Based on the ratio of the effective thermal conductivity of the solid and liquid phases in the inner flow region, the heat flux density at the boundary between the solid and liquid phases of the porous medium in the inner flow region is calculated. Step 5) Taking into account the seepage and phase change process of coolant in the internal flow region of the aircraft surface, as well as the influence of the capillary structure of the porous medium on the liquid phase flow, a non-thermal equilibrium model is adopted to establish a flow heat transfer calculation model inside the aircraft surface structure and solve the transient physical quantity distribution in the internal flow region of the aircraft. Step 6) Repeat steps 3, 4 and 5. During the continuous iteration, the data at the interface between the inner and outer flow regions interact continuously, and the difference between the physical quantities on both sides of the interface gradually narrows. When the five conditions of continuous temperature, continuous heat flux density, continuous pressure, continuous flow rate and continuous composition are achieved on both sides of the interface between the inner and outer flow regions, the iteration process terminates. The distribution of physical quantities in the inner and outer flow regions in the last iteration is the distribution of physical quantities on the surface of the aircraft considering the coupling of the inner and outer flow fields at the current time step. Step 7) Repeat steps 3 to 6 to obtain the distribution of physical quantities on the surface of the aircraft at different time steps, and finally clarify the transient change process of the physical quantities of temperature and heat flux on the surface of the aircraft considering the coupling of internal and external flow fields during the entire service process.

2. The method for simulating transient heat and mass transfer processes on the surface of an aircraft based on partitioned modeling and interface coupling according to claim 1, characterized in that, The calculation model for multi-component fluid flow and heat transfer in the outflow region is as follows: In the formula, ρ is the density of the fluid in the outer flow region, and u is the velocity of the fluid in the outer flow region. i u j u k and u l Let i, j, k, and l represent the components of the fluid velocity in the outflow region along the directions i, j, k, and l, respectively, and x be the coordinates of the fluid in the outflow region. i x j x k and x l denoted by i, j, k, and l respectively, p is the pressure of the fluid in the outer flow region, T is the temperature of the fluid in the outer flow region, h is the specific enthalpy of the fluid in the outer flow region, E is the total energy of the fluid in the outer flow region, μ is the dynamic viscosity coefficient of the fluid in the outer flow region, and t is the flow time. δ ij It is a component of the unit second-order tensor, and its value is 1 if and only if i = j, otherwise it is 0; It is the Reynolds stress, where u′ i and u′ j Let be the fluctuating velocities of the fluid in the i and j directions within the outflow region, respectively; k, ε, and μ are the pulsating velocities of the fluid in the i and j directions, respectively. t These represent the turbulent kinetic energy, dissipation rate, and turbulent viscosity coefficient of the fluid, respectively; σ k and σ ε These are the Prandtl numbers for turbulent kinetic energy and dissipation rate, respectively; C1 and C2 are both constants; Y n D is the mass fraction of the nth component in the fluid in the outflow region. m,n and D T,n These are the mass diffusion coefficient and Soret diffusion coefficient of the nth component in the outflow region, respectively; h n and , respectively, are the specific enthalpy and diffusion flux of the nth component in the outflow region fluid; (τ ij ) eff It is the deviatoric stress tensor of the fluid in the outflow region; S m and S h Representing the mass source term and energy source term in the coolant phase change process, respectively, the Lee model can be used to describe the above source terms as follows: S h =S m h0 In the formula, C evap T is the evaporation constant of the Lee model. sat h0 and k represent the phase change temperature and phase change enthalpy of the coolant, respectively; eff This represents the equivalent thermal conductivity of the fluid in the outflow region; the various physical properties of the fluid in the outflow region can be calculated using the following methods: In the formula, R is the molar gas constant; c p M represents the specific isobaric heat capacity of the fluid in the outflow region; n c p,n k eff,n and μ n These are the relative molar mass, specific heat capacity at constant pressure, equivalent thermal conductivity, and dynamic viscosity coefficient of the nth component in the outflow region fluid, respectively.

3. The method for simulating transient heat and mass transfer processes on the surface of an aircraft based on partitioned modeling and interface coupling according to claim 2, characterized in that, The solution process for the multi-component fluid flow and heat transfer calculation model in the outflow region is as follows: First, set initial conditions. When solving the flow and heat transfer process in the outer flow region of the first time step, use the distribution of physical quantities set in step 2 as the initial conditions; when not solving the first time step, use the distribution of each physical quantity in the outer flow region of the previous time step as the initial conditions for solving this time step. Secondly, boundary conditions are set. The flow rate and temperature of the coolant overflowing from the inner flow region in the calculation results of the previous iteration are used as the boundary conditions of the interface between the inner and outer flow regions. The Mach number of the outer flow and the static pressure and static temperature of the outer flow air are calculated based on the current speed and altitude of the aircraft and used as the boundary conditions for the inlet and outlet of the outer flow region. Furthermore, using the finite volume method, based on the laws of conservation of mass, momentum, and energy, the partial differential equations in the multi-component fluid flow and heat transfer calculation model in the external flow region are transformed into a linear difference equation system, which is then solved. Finally, after the calculation converges, the temperature and pressure distributions of the outflow region in this time step and this iteration step are obtained. Based on the calculation results, the pressure and heat flux density on the outer side of the interface between the inflow and outflow regions are output. The heat flux density is calculated as follows: In the formula, q o T represents the heat flux density on the outer side of the interface between the inflow and outflow regions. 1,o and T 0,o Here, represents the temperature at the center of the boundary grid and the center of the boundary surface, respectively, and ds represents the distance from the center of the boundary grid to the center of the boundary surface. and These are the volume vector of the boundary mesh and the unit normal vector of the boundary surface, respectively.

4. The method for simulating transient heat and mass transfer processes on the surface of an aircraft based on partitioned modeling and interface coupling according to claim 3, characterized in that, The discretization and solution of the partial differential equations in the multi-component fluid flow and heat transfer calculation model in the external flow region are completed using the density-based solver in ANSYS FLUENT software. During the solution process, the physical property settings of the fluid in the external flow region, the loading of flow boundary conditions at the interface between the internal and external flow regions, and the output and saving of the calculation results are achieved through user-defined functions in FLUENT.

5. The method for simulating transient heat and mass transfer processes on the surface of an aircraft based on partitioned modeling and interface coupling according to claim 1, characterized in that, The data interaction method at the interface between the internal and external flow regions in step 4) is implemented based on the AAS module in Matlab software. Through this module, the interaction of temperature, pressure, flow rate, composition and heat flux density on both sides of the interface between the internal and external flow regions is realized, and it is determined whether the physical quantities on both sides of the interface between the internal and external flow regions are continuous, thereby helping to simulate the transient heat and mass transfer process on the surface of the aircraft.

6. The method for simulating transient heat and mass transfer processes on the surface of an aircraft based on partitioned modeling and interface coupling according to claim 1, characterized in that, In step 4), the heat flux density at the boundary between the solid and liquid phases of the porous medium in the internal flow region is expressed as: In the formula, e represents the porosity of the porous medium in the internal region, and k s and k f q represents the thermal conductivity of the solid and liquid phases in a porous medium, respectively. i,f and q i,s q represents the heat flux density at the solid-phase boundary and liquid-phase boundary of the porous medium in the internal flow region, respectively, while q i This represents the total heat flux density at the boundary of the inflow region.

7. The method for simulating transient heat and mass transfer processes on the surface of an aircraft based on partitioned modeling and interface coupling according to claim 1, characterized in that, The heat transfer calculation model inside the surface structure of the aircraft is as follows: In the formula, γ is the velocity of the fluid within the porous medium in the inflow region, K is the permeability of the porous medium in the inflow region, μ is the dynamic viscosity coefficient of the coolant under two-phase mixing conditions, and γ is the velocity of the fluid within the porous medium in the inflow region. H and Γ H ρ represents the coefficients of the convection and diffusion terms in the enthalpy equation, respectively. s ,c p,s ,k s and T s These are the density, specific isobaric heat capacity, thermal conductivity, and temperature of the porous medium solid phase in the inflow region, respectively. sf It refers to the heat exchange between the solid and liquid phases of the porous medium in the inflow region. Let H represent divergence, and H be the enthalpy of the coolant in the internal flow region, expressed as follows: H=ρh-H0,H0=ρh v,sat In the formula, H0 is the coolant enthalpy reference point set in the calculation, h is the specific enthalpy of the coolant, and h v,sat This is the saturated vapor enthalpy of the coolant.

8. The method for simulating transient heat and mass transfer processes on the surface of an aircraft based on partitioned modeling and interface coupling according to claim 7, characterized in that, The solution process for the heat transfer calculation model inside the surface structure of the aircraft is as follows: First, set initial conditions. When solving the flow and heat transfer process in the inflow region of the first time step, use the distribution of physical quantities in the inflow region set in step 2 as the initial conditions; when not solving the first time step, use the distribution of each physical quantity in the inflow region of the previous time step as the initial conditions for solving this time step. Secondly, boundary conditions are set. In the calculation results of the previous iteration, the pressure and heat flux density of the fluid outside the interface between the inner and outer flow fields are used as the boundary conditions at the interface between the inner and outer flow regions. The temperature and set flow rate of the coolant inside the aircraft are used as the boundary conditions for the coolant inlet. Furthermore, the finite volume method is used to transform the partial differential equations in the calculation model of flow and heat transfer inside the surface structure of the aircraft into a linear difference equation system, which is then solved. Finally, after the calculation converges, the temperature and velocity distribution of the internal flow region in this time step and this iteration step are obtained. Based on the calculation results, the composition, temperature, and flow rate of the coolant on the inner side of the interface between the internal and external flow fields are output. The calculation methods for the coolant temperature and flow rate are as follows: T i =eT f +(1-e)T s In the formula, T i It is the temperature on the inner boundary grid of the interface between the inflow and outflow regions, T f and T s q represents the temperature at the center point of the liquid and solid phase boundary grids, respectively. m,i It is the flow rate of coolant overflowing from the internal flow area.

9. The method for simulating transient heat and mass transfer processes on the surface of an aircraft based on partitioned modeling and interface coupling according to claim 8, characterized in that, The discretization and solution of the partial differential equations in the heat transfer calculation model inside the surface structure of the aircraft are completed using the pressure-based solver in ANSYS FLUENT software. During the solution process, the setting of coolant properties, the loading of heat flux density boundary conditions at the interface between the internal and external flow fields, and the output and saving of calculation results are achieved through user-defined functions in FLUENT.

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