CN115659854A - Method for simulating aircraft surface transient heat and mass transmission process based on partition modeling and interface coupling - Google Patents
Method for simulating aircraft surface transient heat and mass transmission process based on partition modeling and interface coupling Download PDFInfo
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Abstract
The invention discloses a method for simulating the transient heat and mass transmission process of the surface of an aircraft based on partition modeling and interface coupling. The invention comprehensively considers the flow of the coolant in the porous shell of the aircraft under the operation condition and the unsteady heat mass transmission process between the coolant and the external high-Mach-number flow field, can accurately predict the temperature and the heat flow distribution of the surface structure of the aircraft under the operation condition, and has guiding significance for the fine design of the surface heat structure and the cooling system of the aircraft.
Description
Technical Field
The invention belongs to the technical field of aircraft thermal management, and particularly relates to a method for simulating a transient thermal mass transmission process of an aircraft surface based on partition modeling and interface coupling.
Background
The high-speed flight technology is regarded as a subversive technology in the future and will have revolutionary influence on the design concept of the manned aerospace vehicle. The speed of the future high-speed aircraft is close to the limit of long-time maneuvering flight in the atmosphere, and the high-speed aircraft has the characteristics of long operation time, high flight speed and the like. However, when such an aircraft flies at a high speed and a long time in a thin atmosphere, the aerodynamic heating is very severe, and a cooling technology must be applied to eliminate the influence of the aerodynamic heat on the internal electronic devices of the aircraft. Therefore, the development of small-space and high-efficiency cooling technology becomes one of the key points in the design and development process of the aircraft. Among conventional cooling techniques, the sweat cooling technique has a high cooling capacity and a small demand for a coolant, and has received a high degree of attention in recent years from the academic world and the industrial world. How to optimally design the aircraft sweating cooling system to accurately match the distribution characteristics of the surface heat load of the aircraft becomes the key for improving the heat protection capability of the aircraft. Before optimizing the design, however, the transient changes in the surface temperature and heat flow of the aircraft under real flight conditions must be captured to clarify the crux of the existing system. In consideration of the complexity of a flight experiment, the method of numerical calculation is adopted to simulate the heat and mass transmission process in the inner and outer flow areas of the aircraft, so that the transient temperature and the heat flow distribution change of the surface of the aircraft are obtained, and the method has higher feasibility. However, a number of challenges remain in the operation of numerical simulation of heat and mass transfer processes in the inflow and outflow regions of an aircraft. Which comprises the following steps: (1) A calculation method for coupling and considering high Mach number flow outside the aircraft and low-speed flow inside the aircraft is lacked; (2) In the operation condition, the speed and the flying height of the aircraft change along with the time, however, a method for researching the influence of the transient change of the external flow field on the internal flow field is lacked at present.
At present, scholars at home and abroad develop a series of researches on the surface heat and mass transmission process of an aircraft considering the influence of an external flow field. And the Dong and Wang adopt a semi-mixed phase model to carry out numerical study on the heat and mass transmission process in the sweating cooling system, and analyze the influence of the thermal conductivity, the porosity and the like of the porous framework on the cooling capacity of the system. According to the article, the influence of an aircraft outer flow field on the flow heat exchange process of an inner flow region is reflected by applying a non-uniform heat flow boundary condition on a heated end of the inner flow region. Ribosch et al developed a multi-physical-quantity coupling system for simulating steady-state sweating cooling, which could be used to analyze the heat and mass transfer processes inside and outside the aircraft structure under steady-state service conditions. Su et al combined a two-phase mixture model to study the heat and mass transfer process in the aircraft transpiration cooling system under steady state conditions, and analyzed the influence of parameters such as porous structure particle diameter on cooling efficiency. From the analysis, it can be seen that, in the research of the surface thermal mass transmission process of the existing aircraft, the simulation of the unsteady high-mach-number external flow field is simplified to different degrees, and the influence of the transient-change aircraft external flow field on the internal flow field is generally difficult to research.
Disclosure of Invention
Aiming at the defects of the existing numerical simulation method of the aircraft surface heat and mass transmission process, the invention provides a method for simulating the aircraft surface transient heat and mass transmission process based on partition modeling and interface coupling. The method can obtain the transient change of the temperature and the heat flow distribution on the surface of the aircraft in the outflow field with time variation, so as to help realize the fine design of the cooling system on the surface of the aircraft.
The method provided by the invention comprises the following steps:
a method for simulating the transient heat and mass transfer process of the surface of an aircraft based on partition modeling and interface coupling is suitable for the aircraft adopting a phase-change sweating cooling system, and comprises the following steps:
step 1: respectively establishing physical models of an inner flow region on the surface of the aircraft and an inner flow region and an outer flow region in the adjacent space according to the real structure of the aircraft, further respectively dividing grids for the inner flow region and the outer flow region, realizing discretization of a physical space, and performing independence check on the grid division of the inner flow region and the outer flow region according to the convergence of heat flow and pressure on interfaces of the inner flow region and the outer flow region;
and 2, step: determining the static temperature and the static pressure of the outflow air as initial conditions of the outflow region by combining the actual operation working conditions of the aircraft, and determining the initial temperature, pressure and flow distribution of the interface of the inner and outer flow regions and the inner flow region according to the temperature, pressure and flow of the fluid in the inner flow region;
and step 3: comprehensively considering the flowing, diffusing and phase-changing processes of the coolant overflowing from the inner flow region in the outer flow region and the compressible property of the air in the outer flow region, describing the turbulent flow in the outer flow region by adopting a k-epsilon model, and establishing a multi-component fluid flowing heat transfer calculation model in the outer flow region as follows:
where ρ is the density of the fluid in the outflow region, u is the velocity of the fluid in the outflow region, and u is i 、u j 、u k And u l Respectively representing the fluid velocity of the outflow regionThe components of the degree in the i, j, k and l directions, x is the coordinate of the fluid in the outflow region, x i 、x j 、x k And x l Respectively representing the components of the fluid coordinates in the directions i, j, k and l, wherein p is the pressure of the fluid in the outflow region, T is the temperature of the fluid in the outflow region, h is the specific enthalpy of the fluid in the outflow region, E is the total energy of the fluid in the outflow region, mu is the dynamic viscosity coefficient of the fluid in the outflow region, and T is the flow time; delta ij Is a component of a unit second order tensor, whose value is 1 if and only if i = j, and 0 otherwise;
is Reynolds stress, of which u' i And u' j The pulsation speeds of the fluid in the direction i and the direction j in the outflow area are respectively; k, ε and μ t Turbulent pulsating energy, dissipation ratio and turbulent viscosity coefficient of the fluid are respectively; sigma k And σ ε Turbulent prandtl numbers which are turbulent pulsation energy and dissipation ratio, respectively; c 1 And C 2 Are all constants; y is n Is the mass fraction of the nth component in the fluid in the outflow region, D m,n And D T,n Mass diffusion coefficient and Soret diffusion coefficient of the nth component in the fluid of the outflow region respectively; h is n And
specific enthalpy and diffusion flux of the nth component in the fluid in the outflow region respectively; (τ) ij ) eff Is the bias stress tensor of the fluid in the outflow region; s. the m And S h Respectively representing a mass source item and an energy source item in the phase change process of the coolant, wherein the source items are described by adopting a Lee model to obtain:
S h ＝S m ·h 0
in the formula, C evap Evaporation constant, T, for Lee model sat And h 0 Respectively representing the phase change temperature and the phase change enthalpy of the coolant; k is a radical of formula eff Representing the equivalent thermal conductivity of the fluid in the outflow region; the physical properties of the fluid in the outflow region can be calculated by the following methods:
wherein R is a molar gas constant; c. C p Showing the specific constant pressure heat capacity of the fluid in the outflow region; m n 、c p,n 、k eff,n And mu n The relative molar mass, specific constant pressure heat capacity, equivalent thermal conductivity and kinetic viscosity coefficient of the nth component in the fluid in the outflow region, respectively.
Firstly, setting initial conditions, and taking the physical quantity distribution set in the step 2 as the initial conditions when solving the flowing heat transfer process in the outer flow region at the first time step; when the first time step is not solved, the distribution of each physical quantity in the area outside the time step is used as the initial condition for solving the time step. Secondly, setting boundary conditions, wherein in the calculation result of the iteration step, the flow and the temperature of the coolant overflowing from the inner flow area are used as boundary conditions of the interface of the inner flow area and the outer flow area; and calculating the outer flow Mach number and the static pressure and the static temperature of the outer flow air as boundary conditions of the inlet and the outlet of the outer flow area according to the speed and the altitude of the current aircraft. And thirdly, converting a partial differential equation set in the multi-component fluid flow heat transfer calculation model in the outflow region into a linear differential equation set by adopting a finite volume method based on mass conservation, momentum conservation and energy conservation laws, and solving. And finally, obtaining the temperature and pressure distribution of the time step and the outflow region in the iteration step after the calculation is converged. And outputting the pressure and the heat flow density outside the interface of the inner flow region and the outer flow region according to the calculation result, wherein the calculation method of the heat flow density comprises the following steps:
in the formula, q o For the heat flow density, T, outside the interface of the inner and outer flow regions 1,o And T 0,o The temperature at the center point of the boundary mesh and the temperature at the center point of the boundary face, respectively, ds represents the distance from the center of the boundary mesh to the center of the boundary face,
and
the volume vector of the boundary mesh and the unit normal vector of the boundary surface are respectively.
And 4, step 4: designing an interaction method of data on the interface of the inner flow area and the outer flow area according to the calculation result of the outer flow area; and setting the pressure and the heat flow density of the fluid outside the interface of the inner flow area and the outer flow area as boundary conditions for calculating the flow heat transfer of the inner flow area. The aircraft surface inflow region is composed of a porous medium, and the thermal diffusion coefficient of a solid phase structure in the porous medium is far higher than that of a coolant, so that under the condition of external heat flow, the temperature and the heat flow of a solid phase and a liquid phase on the interface of the internal and external flow regions are greatly different. According to the ratio of the effective thermal conductivities of the solid phase and the liquid phase, the heat flow density of the solid phase and the liquid phase boundary of the porous medium in the inflow region can be expressed as follows:
wherein e is the porosity of the porous medium in the inner region, k s And k f Thermal conductivity of solid and liquid phases in porous media, q, respectively i,f And q is i,s Represents the heat flow density of the solid phase boundary and the liquid phase boundary of the porous medium in the inflow region respectively, and q i Representing the total heat flux density at the boundary of the inner flow region.
And 5: comprehensively considering the phase change process of the coolant in the flow area in the surface of the aircraft and the influence of the capillary structure of the porous medium on the flow of the liquid phase, a non-thermal equilibrium model (LTNE) is adopted to combine a Two-phase mixed model (TPMM) and a Two-phase mixed model, and the flow heat transfer calculation model in the surface structure of the aircraft is established as follows:
in the formula (I), the compound is shown in the specification,
is the velocity of the fluid in the porous medium in the inner flow region, K is the permeability of the porous medium in the inner flow region, μ is the kinetic viscosity coefficient of the coolant under two-phase mixing conditions, γ H And Γ H Coefficient, p, representing the convective phase and diffusion terms in the enthalpy equation, respectively s ,c p,s ,k s And T s Density, specific constant pressure heat capacity, thermal conductivity and temperature, Q, of the porous medium solid phase in the inner flow region sf Is the heat exchange quantity between the solid phase and the liquid phase of the porous medium in the inner flow area,
expressing divergence, H is the enthalpy of the inner flow region coolant, expressed as follows:
H＝ρh-H 0 ,H 0 ＝ρh v,sat
in the formula, H 0 For the coolant enthalpy reference point set in the calculation, h is the specific enthalpy of the coolant, h v,sat Is the saturated steam specific enthalpy of the coolant.
Firstly, setting initial conditions, and taking the physical quantity distribution of the internal flow region set in the step 2 as the initial conditions when solving the flow heat transfer process of the internal flow region in the first time step; when the first time step is not solved, the distribution of each physical quantity of the flow region in the previous time step is used as the initial condition for solving the time step. Secondly, setting boundary conditions, and taking the pressure and the heat flow density of the fluid outside the interface of the inner flow field and the outer flow field as the boundary conditions at the interface of the inner flow field and the outer flow field in the calculation result of the iteration step. The temperature and the set flow of the coolant in the aircraft are used as boundary conditions for the coolant inlet. And thirdly, converting a partial differential equation set in the flow heat transfer calculation model in the aircraft surface structure into a linear differential equation set by adopting a finite volume method, and solving. And finally, after the calculation is converged, the temperature and flow velocity distribution of the time step and the inner flow region in the iteration step is obtained. Outputting the components, the temperature and the flow of the coolant inside the interface of the inner flow field and the outer flow field according to the calculation result, wherein the calculation methods of the temperature and the flow of the coolant respectively comprise the following steps:
T i ＝eT f +(1-e)T s
in the formula, T i Is the temperature, T, on the inside boundary grid of the interface of the inside and outside flow regions f And T s Temperature at the center of the boundary grid of the liquid and solid phases, q m,i Is the flow rate of coolant spilled in the inner flow region.
Step 6: and (5) repeating the step (3), the step (4) and the step (5), wherein in the continuous iteration process, data on the interface of the internal and external flow areas are continuously interacted, and the difference between the physical quantities on the two sides of the interface is gradually reduced. When the two sides of the interface of the inner and outer flow areas realize five conditions of continuous temperature, continuous heat flow density, continuous pressure, continuous flow and continuous components, the iteration process is terminated, and the physical quantity distribution of the inner and outer flow areas in the last iteration is the surface physical quantity distribution of the aircraft considering the coupling of the inner and outer flow areas under the current time step.
And 7: and (5) repeating the steps 3 to 6, obtaining the physical quantity distribution of the aircraft surface under different time steps, and finally determining the transient change process of the physical quantities such as the aircraft surface temperature and the heat flow coupled by the internal and external flow fields in the full service process.
Compared with the prior art, the invention has the beneficial effects that:
1) Based on the thought of 'partition modeling and interface coupling', the high Mach number flow of the air outside the aircraft and the low-speed flow of the coolant inside the aircraft are separately modeled and solved, and a more appropriate solver and a discrete method for pressure-speed coupling are distributed according to the physical characteristics of the two regions; an interface coupling algorithm is introduced, continuity of temperature, pressure, flow, components and heat flow density of the two regions on the boundary is achieved, and the transient change process of the surface temperature and heat flow distribution of the aircraft under the influence of an external flow field is defined.
2) The influence of the time-varying property of the external flow field of the aircraft on the heat and mass transmission process of the internal flow field of the aircraft in the service process is determined; the calculation result can guide the fine design of the aircraft surface cooling system, and the real-time accurate matching between the coolant flow and the external transient heat flow load is realized, so that the aims of high efficiency and light weight of the cooling system are fulfilled.
Drawings
FIG. 1 is a numerical method for simulating the transient heat and mass transfer process of an aircraft surface based on the partitioned modeling and interface coupling of the invention;
FIG. 2 is a physical model of a surface-sweating cooled part of an aircraft under supersonic main flow with an angle of attack of 20 degrees and a Mach number of 4.5
FIG. 3 meshing of the external flow field and surface transpiration cooling elements of an aircraft
FIG. 4 verification of grid independence of aircraft external flow field and surface transpiration cooling elements
FIG. 5 is a transient change process of the coolant flow distribution on the surface of an aircraft under supersonic main flow with an attack angle of 20 degrees and a Mach number of 4.5
FIG. 6 is a transient variation process of the distribution of the heat flux density on the surface of the aircraft under supersonic main flow with an attack angle of 20 degrees and a Mach number of 4.5
FIG. 7 is a diagram showing the transient variation process of the temperature distribution on the surface of the aircraft under supersonic main flow with an attack angle of 20 degrees and a Mach number of 4.5
FIG. 8 is a diagram showing a transient change process of the pressure distribution on the surface of an aircraft under supersonic main flow with an attack angle of 20 degrees and a Mach number of 4.5
Detailed Description
The invention is described in detail below by taking a two-dimensional sweating cooling system under a supersonic flow field as an example with reference to the accompanying drawings:
as shown in FIG. 1, the method for simulating the transient heat and mass transfer process of the aircraft surface based on the partition modeling and the interface coupling of the invention comprises the following steps:
first, an aircraft surface slab structure and its onboard phase change transpiration cooling system are shown in fig. 2. The aircraft surface structure is composed of a metal structural part and a porous medium area, and a water storage tank is arranged below the porous medium area. Under the working condition of service, the supersonic main flow outside the aircraft bypasses the surface of the flat plate and forms a shock wave surface, the temperature of the air flow behind the shock wave surface can be increased violently, at the moment, the coolant in the water storage tank generates gas-liquid phase change and flows through the porous medium area to the outflow area, and the sweating cooling of the surface of the aircraft is realized. Due to the morphological characteristics of the aircraft surface flat plate structure, the three-dimensional structure can be equivalent to a two-dimensional plane model so as to reduce the demand of computing resources. Physical models of an inner flow area and an ultrasonic outer flow area of the flat sweat cooling system are respectively established, structured grids are respectively divided into the inner flow area and the outer flow area, the grids of the near-interface parts in the inner flow area and the outer flow area are encrypted, and the grid division results are shown in fig. 3. Further, the mesh of the inner and outer flow regions was subjected to the irrelevance check based on the convergence of the heat flow and the pressure at the interface between the inner and outer flow fields, and the result is shown in fig. 4. As can be seen from fig. 4, when the number of the grids is higher than 235800, the influence of the heat flow density and the fluid pressure calculation result on the interface of the internal and external flow regions by the continuous encryption grid is less than 1%, so that the set of grids is selected to perform simulation calculation on the transient heat mass transmission process of the aircraft flat plate structure surface;
secondly, selecting the static temperature and the static pressure of outflow air to be 226.5K and 1172Pa respectively; selecting water as a coolant in the sweating cooling system, wherein the initial temperature of the coolant is 300K; the flow rate of the coolant at the inlet of the system is set to be 0.3 kg.m -2 ·s -1 (ii) a Setting initial conditions of the inner and outer flow areas according to the working conditions;
thirdly, comprehensively considering the flowing, diffusing and phase-changing processes of the coolant overflowing from the inner flow region in the outer flow region and the compressible property of the air in the outer flow region, describing the turbulent flow of the fluid in the outer flow region by adopting a k-epsilon model, and establishing a multi-component fluid flowing heat transfer calculation model in the outer flow region as follows:
where ρ is the density of the fluid in the outflow region, u is the velocity of the fluid in the outflow region, and u is the velocity of the fluid in the outflow region i 、u j 、u k And u l Representing the components of the fluid velocity in the i, j, k and l directions, respectively, of the outflow region, x being the coordinates of the fluid in the outflow region, x i 、x j 、x k And x l Representing the components of the fluid coordinates in the directions i, j, k and l respectively, p is the pressure of the fluid in the outflow region, T is the temperature of the fluid in the outflow region, h is the specific enthalpy of the fluid in the outflow region, E is the total energy of the fluid in the outflow region, μ is the dynamic viscosity coefficient of the fluid in the outflow region, and T is the flow time; delta ij Is a component of the unit second order tensor, whose value is 1 if and only if i = j, and 0 otherwise;
is Reynolds stress, of which u' i And u' j The pulsation speeds of the fluid in the direction i and the direction j in the outflow area are respectively; k, ε and μ t Turbulent pulsating energy, dissipation ratio and turbulent viscosity coefficient of the fluid are respectively; sigma k And σ ε Turbulence prandtl numbers, which are turbulence pulsation energy and dissipation rate, respectively; c 1 And C 2 Are all constants; y is n Is the mass fraction of the nth component in the fluid in the outflow region, D m,n And D T,n Mass diffusion coefficient and Soret diffusion coefficient of the nth component in the fluid of the outflow region respectively; h is n And
specific enthalpy and diffusion flux of the nth component in the fluid in the outflow region, respectively; (τ) ij ) eff Is the bias stress tensor of the fluid in the outflow region; s. the m And S h Respectively representing mass source terms and energies during phase change of coolantAnd (3) a source item is described by adopting a Lee model, so that the following can be obtained:
S h ＝S m ·h 0
in the formula, C evap Evaporation constant, T, for Lee model sat And h 0 Respectively representing the phase change temperature and the phase change enthalpy of the coolant; k is a radical of eff Representing the equivalent thermal conductivity of the fluid in the outflow region; the physical properties of the fluid in the outflow region can be calculated by the following methods:
wherein R is a molar gas constant; c. C p Indicating the specific constant pressure heat capacity of the fluid in the outflow region; m n 、c p,n 、k eff,n And mu n The relative molar mass, specific constant pressure heat capacity, equivalent thermal conductivity and kinetic viscosity coefficient of the nth component in the fluid in the outflow region are respectively.
Firstly, setting initial conditions, and taking the physical quantity distribution set in the step 2 as the initial conditions when solving the flowing heat transfer process in the outer flow region at the first time step; when the first time step is not solved, the distribution of each physical quantity in the area outside the previous time step is taken as the initial condition for solving the time step. Secondly, setting boundary conditions, wherein in the calculation result of the iteration step, the flow and the temperature of the coolant overflowing from the inner flow area are used as boundary conditions of the interface of the inner flow area and the outer flow area; and calculating the Mach number of the outflow and the static pressure and the static temperature of the outflow air as boundary conditions of the inlet and the outlet of the outflow area according to the speed and the altitude of the current aircraft. And thirdly, converting a partial differential equation set in the multi-component fluid flow heat transfer calculation model in the outflow region into a linear differential equation set by adopting a finite volume method based on mass conservation, momentum conservation and energy conservation laws, and solving. And finally, obtaining the temperature and pressure distribution of the time step and the outflow region in the iteration step after the calculation is converged. And outputting the pressure and the heat flow density outside the interface of the inner region and the outer region according to the calculation result, wherein the calculation method of the heat flow density comprises the following steps:
in the formula, q o For the heat flow density, T, outside the interface of the inner and outer flow regions 1,o And T 0,o Respectively, the temperature at the center point of the boundary grid and the center point of the boundary surface, ds represents the distance from the center of the boundary grid to the center of the boundary surface,
and
the volume vector of the boundary mesh and the unit normal vector of the boundary surface are respectively.
The dispersion and solution of partial differential equations in the multi-component fluid flow heat transfer calculation model in the outflow region can be completed by means of a density basis solver in ANSYS FLUENT software, and in the solution process, the physical property setting of the fluid in the outflow region, the loading of flow boundary conditions on an interface of the inflow region and the outflow region, and the output and storage of calculation results are realized through a User Defined Function (UDF) in FLUENT.
Fourthly, acquiring the boundary condition of the flow heat transfer calculation in the inner flow area according to the pressure and the heat flow density of the fluid at the interface of the inner flow area and the outer flow area in the calculation result of the outer flow area. The aircraft surface inflow region is composed of a porous medium, and the thermal diffusion coefficient of a solid phase structure in the porous medium is far higher than that of a coolant, so that under the condition of external heat flow, the temperature and the heat flow of a solid phase and a liquid phase on the interface of the internal and external flow regions are greatly different. According to the ratio of the effective thermal conductivities of the solid phase and the liquid phase, the heat flow density of the solid phase and the liquid phase boundary of the porous medium in the inflow region can be expressed as follows:
wherein e is the porosity of the porous medium in the inner region, k s And k f Thermal conductivity of solid and liquid phases in porous media, q i,f And q is i,s Represents the heat flow density of the solid phase boundary and the liquid phase boundary of the porous medium in the inflow region respectively, and q i Representing the total heat flux density at the boundary of the inner flow region.
In this step, the method for interacting data on the interface between the internal and external flow areas can be realized based on an AAS (As a Server) module in Matlab software, and by using the module, the interaction of the temperature, the pressure, the flow, the components and the heat flow density at two sides of the interface between the internal and external flow areas can be realized, and whether the physical quantities at two sides of the interface between the internal and external flow areas are continuous or not can be judged, thereby helping to realize the simulation of the transient heat and mass transmission process of the surface of the aircraft.
Fifthly, comprehensively considering the seepage and phase change processes of the coolant in the flow area of the aircraft surface and the influence of the capillary structure of the porous medium on the liquid phase flow, and establishing a flow heat transfer calculation model inside the aircraft surface structure by adopting a non-thermal equilibrium model and a two-phase mixing model as follows:
in the formula (I), the compound is shown in the specification,
is the velocity of the fluid in the porous medium in the inner flow region, K is the permeability of the porous medium in the inner flow region, μ is the kinetic viscosity coefficient of the coolant under two-phase mixing conditions, γ H And Γ H Coefficient, p, representing respectively the convective phase and diffusion terms in the enthalpy equation s ,c p,s ,k s And T s Density, specific constant pressure heat capacity, thermal conductivity and temperature, Q, of the porous medium solid phase in the inner flow region sf Is the heat exchange quantity between the solid phase and the liquid phase of the porous medium in the inner flow area,
expressing divergence, H is the enthalpy of the inner flow region coolant, expressed as follows:
H＝ρh-H 0 ,H 0 ＝ρh v,sat
in the formula, H 0 For the coolant enthalpy reference point set in the calculation, h is the specific enthalpy of the coolant, h v,sat Is the saturated steam specific enthalpy of the coolant.
Firstly, setting initial conditions, and taking the physical quantity distribution set in the step 2 as the initial conditions when solving the flow heat transfer process of the internal flow region in the first time step; when the first time step is not solved, the distribution of each physical quantity of the flow region in the previous time step is used as the initial condition for solving the time step. Secondly, setting boundary conditions, namely taking the pressure and the heat flow density of the fluid outside the interface between the inner flow field and the outer flow field as the boundary conditions in the calculation result of the iteration step. And thirdly, converting a partial differential equation set in the flow heat transfer calculation model in the aircraft surface structure into a linear differential equation set by adopting a finite volume method, and solving. And finally, obtaining the temperature and flow velocity distribution of the time step and the inner flow region in the iteration step after the calculation is converged. Outputting the components, the temperature and the flow of the coolant at the inner side of the interface of the inner flow field and the outer flow field according to the calculation result, wherein the calculation methods of the temperature and the flow of the coolant respectively comprise the following steps:
T i ＝eT f +(1-e)T s
in the formula, T i Is the temperature, T, on the inside boundary grid of the interface of the inside and outside flow regions f And T s Temperature at the center of the boundary grid of the liquid and solid phases, q m,i Is the flow rate of the overflow of the coolant in the inner flow region.
Similarly, the dispersion and solution of partial differential equations in the flow heat transfer calculation model in the aircraft surface structure can be completed by means of a pressure-based solver in ANSYS FLUENT software, and in the solution process, the physical property setting of the coolant, the loading of the heat flow density boundary conditions at the interface of the inner flow field and the outer flow field, and the output and storage of the calculation result can be realized by user-defined functions in FLUENT.
Sixthly, in the continuous iteration process, the difference between the physical quantities on the two sides of the interface of the internal and external flow areas is gradually reduced. When five conditions of continuous temperature, continuous heat flow density, continuous pressure, continuous flow and continuous components are realized on two sides of the interface of the internal and external flow fields, the iteration process is terminated, and the physical quantity distribution of the internal and external flow fields in the last iteration is the aircraft surface physical quantity distribution considering the coupling of the internal and external flow fields under the current time step.
Seventhly, calculating the transient change of the distribution of the physical quantity on the surface of the flat plate sweating cooling system in 12s time of coupling of the internal flow field and the external flow field, wherein the distribution of the flow, the heat flow, the temperature and the pressure on the surface of the system is respectively shown as the graph 5, the graph 6, the graph 7 and the graphShown in fig. 8. From the transient change of the aircraft surface temperature distribution, the aircraft surface temperature does not exceed 378K at most within 12s, and after the internal field and the external field are contacted for 6s, the amplitude of the aircraft surface maximum temperature is increased by less than 1.1 percent along with the continuous extension of the running time. The above results show that the sweating cooling system using water as working medium can meet the cooling requirement of the aircraft flat plate structure under 4.5 Mach. However, from transient distributions of fluid pressure, flow rate and heat flow density on the surface of the porous medium, when the internal and external fields are contacted for 6s, the coolant on the near supersonic inflow side in the porous medium is completely gasified, and the local coolant flow rate is as low as 0.024kg m -2 ·s -1 Resulting in a reduction in the overall performance of the sweating cooling system. Such a heat transfer deterioration phenomenon occurs because the pressure in the region near the incoming flow in the porous medium rises sharply due to the vaporization of the coolant, resulting in the obstruction of the inflow of the coolant. Under the influence, the coolant flows to the areas, far away from the incoming flow, with low heat flow density and low pressure, in the porous medium, so that the phenomenon of uneven distribution of the coolant is generated in the sweating cooling system. In order to further improve the efficiency of the sweating cooling system of the aircraft flat plate structure, the simulation result is combined, the matching design of the coolant flow transient distribution is provided according to the transient distribution of the heat flow on the surface of the porous medium, namely, a coolant supply strategy which changes along with the air parameter is designed according to the heat flow distribution provided by the simulation, when the heat transfer deterioration phenomenon of the sweating cooling system occurs (9 s after operation), the coolant flow in the area with higher heat flow density is improved, so that the uniform distribution of the coolant flow in the porous medium under the pneumatic heating condition is maintained, and the influence of the heat transfer deterioration phenomenon in the porous medium on the overall performance of the sweating cooling system is eliminated in advance. Different from the traditional coolant supply strategy of the sweating cooling system changing along with the space, the coolant supply strategy changing along with the space parameters based on the design of the patent can realize the matching of the unsteady heat flow of the aircraft, can properly reduce the supply of the coolant when the heat flow density is lower (such as in operation for 9 s), and can also meet the cooling requirement of the surface structure of the aircraft.
The invention provides a numerical method for simulating the transient heat and mass transmission process of the surface of the aircraft based on partition modeling and interface coupling, which can accurately obtain the transient temperature and heat flow density change in the surface sweating cooling system of the aircraft in the real service process and determine the time-space non-uniformity characteristic of the surface heat load distribution of the aircraft under the pneumatic heating condition. On the basis, a coolant flow distribution strategy matched with the surface heat load distribution characteristic can be reasonably designed, the utilization efficiency of the coolant in the sweating cooling system is improved, the requirement of the system on the total amount of the coolant is reduced, and the light design goal of the sweating cooling system is realized. According to the method, the time and the position of the heat transfer deterioration phenomenon in the sweating cooling system can be accurately predicted, and the method has guiding significance for the efficient and light-weight design of the aircraft surface cooling system.
Based on the thought of 'partition modeling and interface coupling', the method separately solves the high Mach number flow of an outer flow area and the low-speed flow of an inner flow area of the aircraft, and designs a transmission method of interface data of the inner flow area and the outer flow area, thereby realizing the transient coupling solution of the inner flow area and the outer flow area of the aircraft. The method can accurately predict the transient temperature and heat flow distribution of the aircraft surface structure under the real outflow condition under the operation working condition, and has guiding significance for the fine design of the aircraft surface heat structure and a cooling system.
Claims (9)
1. A method for simulating the transient heat and mass transmission process of the surface of an aircraft based on partition modeling and interface coupling is suitable for the aircraft adopting a phase-change sweating cooling system, an inner flow area on the surface of the aircraft is composed of a porous medium, and the heat diffusion coefficient of a solid phase structure in the porous medium is far higher than that of a coolant, and is characterized by comprising the following steps of:
step 1) respectively establishing physical models of an inner flow area on the surface of the aircraft and inner and outer flow areas in a neighboring space according to the real structure of the aircraft, further respectively dividing grids of the inner and outer flow areas to realize discretization of a physical space, and performing independence check on the grid division of the inner and outer flow areas according to the convergence of heat flow and pressure on interfaces of the inner and outer flow areas;
step 2) determining the static temperature and the static pressure of outflow air as initial conditions of an outflow region by combining the actual operation working conditions of the aircraft, and determining the initial temperature, pressure and flow distribution of an inner and outer region interface and an inner flow region according to the temperature, pressure and flow of fluid in the inner flow region;
step 3) comprehensively considering the flowing, diffusing and phase-changing processes of the coolant overflowing from the inner flow region in the outer flow region and the compressible property of the air in the outer flow region, describing turbulence in the outer flow region by adopting a k-epsilon model, establishing a multi-component fluid flowing heat transfer calculation model in the outer flow region, and solving the transient physical quantity distribution of the aircraft outer flow region;
step 4) designing an interactive method of data on the interface of the inner flow area and the outer flow area according to the calculation result of the outer flow area; setting the pressure and the heat flow density of the fluid outside the interface surface of the internal and external flow areas as boundary conditions for calculating the flow heat transfer of the internal flow area, and calculating the heat flow density of the boundary of the solid phase and the liquid phase of the porous medium of the internal flow area according to the effective heat conductivity ratio of the solid phase and the liquid phase of the internal flow area;
step 5) comprehensively considering the seepage and phase change processes of the coolant in the flow area of the surface of the aircraft and the influence of the capillary structure of the porous medium on the flow of the liquid phase, establishing a flow heat transfer calculation model inside the surface structure of the aircraft by adopting a non-thermal balance model, and solving the transient physical quantity distribution of the flow area of the aircraft;
and 6) repeating the step 3, the step 4 and the step 5, wherein in the continuous iteration process, data on the interface of the internal and external flow areas are continuously interacted, and the difference between the physical quantities on the two sides of the interface is gradually reduced. When five conditions of continuous temperature, continuous heat flow density, continuous pressure, continuous flow and continuous components are realized on two sides of an interface of the internal and external flow areas, the iteration process is terminated, and the physical quantity distribution of the internal and external flow areas in the last iteration is the surface physical quantity distribution of the aircraft considering the coupling of the internal and external flow areas under the current time step;
and 7) repeating the steps 3 to 6, obtaining the physical quantity distribution of the aircraft surface under different time steps, and finally determining the transient change process of the aircraft surface temperature, heat flow and other physical quantities in consideration of the coupling of the internal flow field and the external flow field in the full service process.
2. The method for simulating an aircraft surface transient heat and mass transport process based on zonal modeling and interface coupling of claim 1, wherein the computational model of multi-component fluid flow heat transfer in the outflow region is represented as follows:
where ρ is the density of the fluid in the outflow region, u is the velocity of the fluid in the outflow region, and u is i 、u j 、u k And u l Representing the components of the fluid velocity in the i, j, k and l directions, respectively, of the outflow region, x being the coordinates of the fluid in the outflow region, x i 、x j 、x k And x l Representing the components of the fluid coordinates in the directions i, j, k and l respectively, p is the pressure of the fluid in the outflow region, T is the temperature of the fluid in the outflow region, h is the specific enthalpy of the fluid in the outflow region, E is the total energy of the fluid in the outflow region, μ is the dynamic viscosity coefficient of the fluid in the outflow region, and T is the flow time; delta ij Is a component of the unit second order tensor, whose value is 1 if and only if i = j, and 0 otherwise;
is Reynolds stress of u' i And u' j The pulsating speeds of the fluid in the outflow region in the directions i and j respectively; k, ε and μ t Turbulent pulsating energy, dissipation ratio and turbulent viscosity coefficient of the fluid are respectively; sigma k And σ ε Turbulent prandtl numbers which are turbulent pulsation energy and dissipation ratio, respectively; c 1 And C 2 Are all constants; y is n Is the mass fraction of the nth component in the fluid in the outflow region, D m,n And D T,n Mass diffusion coefficient and Soret diffusion coefficient of the nth component in the fluid of the outflow region respectively; h is a total of n And
specific enthalpy and diffusion flux of the nth component in the fluid in the outflow region, respectively; (τ) ij ) eff Is the bias stress tensor of the fluid in the outflow region; s. the m And S h Respectively representing a mass source item and an energy source item in the phase change process of the coolant, wherein the source items are described by adopting a Lee model to obtain:
S h ＝S m ·h 0
in the formula, C evap Evaporation constant, T, for Lee model sat And h 0 Respectively representing the phase change temperature and the phase change enthalpy of the coolant; k is a radical of eff Representing the equivalent thermal conductivity of the fluid in the outflow region; outflow zone flowThe physical properties of the body can be calculated by the following methods:
wherein R is a molar gas constant; c. C p Indicating the specific constant pressure heat capacity of the fluid in the outflow region; m n 、c p,n 、k eff,n And mu n The relative molar mass, specific constant pressure heat capacity, equivalent thermal conductivity and kinetic viscosity coefficient of the nth component in the fluid in the outflow region, respectively.
3. The method for simulating an aircraft surface transient heat and mass transfer process based on compartmental modeling and interface coupling of claim 2, wherein the solution of the computational model for multi-component fluid flow heat transfer in the outflow region is as follows:
firstly, setting initial conditions, and taking the physical quantity distribution set in the step 2 as the initial conditions when solving the flow heat transfer process in the first time step outflow region; when the first time step is not solved, the distribution of each physical quantity in the outflow region of the time step is used as the initial condition for solving the time step;
secondly, setting boundary conditions, wherein in the calculation result of the iteration step, the flow and the temperature of the coolant overflowing from the inner flow area are used as boundary conditions of interfaces of the inner flow area and the outer flow area; calculating the Mach number of outflow and the static pressure and static temperature of outflow air as boundary conditions of an inlet and an outlet of an outflow region according to the speed and the altitude of the current aircraft;
thirdly, converting a partial differential equation set in the multi-component fluid flow heat transfer calculation model in the outflow region into a linear differential equation set by adopting a finite volume method based on mass conservation, momentum conservation and energy conservation laws, and solving;
and finally, after the calculation is converged, obtaining the temperature and pressure distribution of the inner and outer flow regions in the time step and the iteration step, and outputting the pressure and heat flow density outside the interface of the inner and outer flow regions according to the calculation result, wherein the calculation method of the heat flow density comprises the following steps:
in the formula, q o For the heat flow density, T, outside the interface of the inner and outer flow regions 1,o And T 0,o The temperature at the center point of the boundary mesh and the temperature at the center point of the boundary face, respectively, ds represents the distance from the center of the boundary mesh to the center of the boundary face,
and
the volume vector of the boundary mesh and the unit normal vector of the boundary surface are respectively.
4. The method for simulating an aircraft surface transient heat and mass transfer process based on partition modeling and interface coupling of claim 3, wherein the discretization and solving of the partial differential equations in the computational model of multi-component fluid flow heat transfer in the outflow region is performed by a density-based solver in ANSYS FLUENT software, and during the solving, the physical property setting of the fluid in the outflow region, the loading of the flow boundary conditions at the interface between the inner and outer flow regions, and the outputting and storing of the calculation results are performed by user-defined functions in FLUENT.
5. The method for simulating the transient heat and mass transfer process on the surface of the aircraft based on the partition modeling and the interface coupling as claimed in claim 1, wherein the interaction method of the data on the interface between the internal and external flow areas in the step 4) is realized based on an AAS module in Matlab software, and by the module, the interaction of the temperature, the pressure, the flow, the components and the heat flow density on the two sides of the interface between the internal and external flow areas is realized, and whether the physical quantities on the two sides of the interface between the internal and external flow areas are continuous or not is judged, thereby helping to realize the simulation of the transient heat and mass transfer process on the surface of the aircraft.
6. The method for simulating the transient heat and mass transfer process of the surface of the aircraft based on the partition modeling and the interface coupling is characterized in that the heat flow density of the solid-phase and liquid-phase boundary of the porous medium in the inner flow region in the step 4) is expressed as follows:
wherein e is the porosity of the porous medium in the inner region, k s And k f Thermal conductivity of solid and liquid phases in porous media, q i,f And q is i,s Represents the heat flow density of the solid phase boundary and the liquid phase boundary of the porous medium in the inflow region respectively, and q i Representing the total heat flux density at the boundary of the inner flow region.
7. The method for simulating the transient heat and mass transfer process on the surface of the aircraft based on the partition modeling and the interface coupling as recited in claim 1, wherein the computational model of the flow heat transfer inside the surface structure of the aircraft is represented as follows:
in the formula (I), the compound is shown in the specification,
is the velocity of the fluid in the porous medium in the inner flow region, K is the permeability of the porous medium in the inner flow region, μ is the kinetic viscosity coefficient of the coolant under two-phase mixing conditions, γ H And gamma H Coefficient, p, representing the convective phase and diffusion terms in the enthalpy equation, respectively s ,c p,s ,k s And T s Density, specific constant pressure heat capacity, thermal conductivity and temperature, Q, of the porous medium solid phase in the inner flow region sf Is the heat exchange quantity between the solid phase and the liquid phase of the porous medium in the inner flow area,
expressing divergence, H is the enthalpy of the inner flow region coolant, which is expressed as follows:
H＝ρh-H 0 ,H 0 ＝ρh v,sat
in the formula, H 0 For the coolant enthalpy reference point set in the calculation, h is the specific enthalpy of the coolant, h v,sat Is the saturated steam specific enthalpy of the coolant.
8. The method for simulating the transient heat and mass transfer process on the surface of the aircraft based on the partition modeling and the interface coupling as claimed in claim 7, wherein the solution process of the flow heat transfer calculation model inside the aircraft surface structure is as follows:
firstly, setting initial conditions, and taking the distribution of the physical quantity of the internal flow region set in the step 2 as the initial conditions when solving the flow heat transfer process of the internal flow region in the first time step; when the first time step is not solved, the distribution of each physical quantity of the flow region in the previous time step is used as the initial condition for solving the time step;
secondly, setting boundary conditions, wherein in the calculation result of the iteration step, the pressure and the heat flow density of the fluid outside the interface of the inner flow field and the outer flow field are used as the boundary conditions of the interface of the inner flow field and the outer flow field, and the temperature and the set flow of the coolant in the aircraft are used as the boundary conditions of a coolant inlet;
thirdly, converting a partial differential equation set in a flow heat transfer calculation model in the aircraft surface structure into a linear differential equation set by adopting a finite volume method, and solving;
and finally, after the calculation is converged, obtaining the temperature and flow velocity distribution of the internal flow region in the time step and the iteration step, and outputting the components, the temperature and the flow of the coolant at the inner side of the interface of the internal flow field and the external flow field according to the calculation result, wherein the calculation methods of the temperature and the flow of the coolant respectively comprise the following steps:
T i ＝eT f +(1-e)T s
in the formula, T i Is the temperature, T, on the inside boundary grid of the interface of the inside and outside flow regions f And T s Temperature at the center of the boundary grid of the liquid and solid phases, q m,i Is the flow rate of the overflow of the coolant in the inner flow region.
9. The method for simulating the transient heat and mass transfer process on the surface of the aircraft based on the partition modeling and the interface coupling is characterized in that the dispersion and the solution of the partial differential equation set in the flow heat transfer calculation model in the aircraft surface structure are completed by a pressure-based solver in ANSYS FLUENT software, and in the solution process, the physical property setting of the coolant, the loading of the boundary condition of the heat flow density at the interface of the inner flow field and the outer flow field, and the output and the storage of the calculation result are realized by a user-defined function in FLUENT.
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