CN115081108B - Hypersonic cruise aircraft full flow field numerical simulation method and system - Google Patents
Hypersonic cruise aircraft full flow field numerical simulation method and system Download PDFInfo
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Abstract
The invention discloses a hypersonic cruise aircraft full flow field numerical simulation method and system, wherein a projectile body and near-field tail flame integrated geometric model and a far-field tail flame geometric model are firstly established, non-structural meshing is conducted on the projectile body and near-field tail flame integrated geometric model, numerical simulation operation is conducted on the projectile body and near-field tail flame integrated geometric model after meshing to obtain a projectile body and near-field tail flame flow field integrated numerical simulation result, structural meshing is conducted on the far-field tail flame geometric model, and numerical simulation operation is conducted on the far-field tail flame geometric model after meshing to obtain a far-field tail flame flow field numerical simulation result. According to the invention, through the integral simulation of the projectile body and the near-field tail flame flow field and the recursive simulation of the far-field tail flame interpolation, the precise simulation of the projectile body and the tail flame flow field of the hypersonic cruise aircraft is realized, and data support can be provided for research of infrared radiation characteristics, aerodynamic effects and the like of a target.
Description
Technical Field
The invention relates to the technical field of hypersonic cruise aircraft flow field simulation, in particular to a hypersonic cruise aircraft full flow field numerical simulation method and system.
Background
The hypersonic cruise aircraft is powered by a scramjet engine, cruises at high speed in the near space with the height of about 30km, has the flying speed exceeding 5Ma, and has the characteristics of high speed, strong maneuverability and the like. The aerodynamic heating effect of the hypersonic cruise aircraft during high-speed maneuvering in the near space can lead to rapid temperature rise of the projectile body, and the Gao Wenwei flame ejected by the scramjet engine is used as a strong infrared radiation source, so that conditions are provided for infrared remote sensing of the hypersonic cruise aircraft. The mastering of the infrared radiation characteristic is the basis of research on infrared remote sensing research, and the calculation of infrared radiation requires the accurate calculation of the flow field distribution of the hypersonic cruise aircraft projectile body and the tail flame. In view of the above, it is significant to accurately calculate the hypersonic cruise aircraft full flow field distribution.
The hypersonic cruise aircraft full flow field mainly comprises an elastomer flow field and a tail flame flow field, and the tail flame is generally divided into a near-field tail flame and a far-field tail flame due to the long and narrow tail flame area. The hypersonic cruise aircraft flow field numerical simulation mainly utilizes a numerical simulation method to calculate the temperature, pressure and other distribution of the projectile surface and surrounding gas, and calculate the parameter distribution rules of temperature, radiant gas content, pressure and other parameters in the tail flame. The hypersonic cruise aircraft projectile and tail flame flow field distribution calculation relates to the theories of computational fluid mechanics, aerodynamic thermodynamics, high-temperature gas jet dynamics, infrared radiation and the like, and the calculation is very complex.
The patent document with the application number 20191012884. X discloses a quick rendering method of infrared radiation characteristics of a hypersonic aircraft, the patent document with the application number 202010205506.1 discloses a high-precision flow-solid coupling calculation method of a nose cone thermal environment of the hypersonic aircraft, the patent document with the application number 201810007755.2 discloses a dynamic temperature modeling method of a hypersonic target surface, and the three published patents only calculate the distribution of an elastomer flow field, do not provide a tail flame flow field calculation method and do not consider the coupling influence of the tail flame flow field and the elastomer flow field.
Qi Junkai, dong Shikui (tactical missile technique, 2021 (5): 17-28) uses numerical simulation methods to simulate the near field flow fields of the projectile and the tail flame respectively, and does not consider the coupling effect of the projectile and the tail flame, and does not provide a calculation method of far-field tail flame.
Zhangxian discloses a master paper-hypersonic aircraft tail flame infrared radiation characteristic study, wherein hypersonic cruise aircraft tail flame flow field distribution is simulated by using a numerical simulation technology, and far-field tail flame flow field distribution is simulated and calculated by using a splicing calculation technology, but the literature does not consider the coupling influence effect of an projectile on a tail flame flow field, and when the far-field flow field is calculated, far-field grids are required to be too stiff in the same division as near-field grids. In practice, during numerical simulation, the dividing density of the grid should be related to the gradient of the data change of the physical field, namely, an encryption grid is used in a region with a large gradient, a sparse grid is used in a region with a small gradient, and the range of the far field region of the tail flame is large and the gradient is small, so that the sparse grid is used.
Disclosure of Invention
The invention aims to solve the technical problem of providing a hypersonic cruise aircraft full flow field numerical simulation method and system, which realize accurate simulation of a hypersonic cruise aircraft projectile body and a far-field tail flame flow field through integral simulation of the projectile body and the near-field tail flame flow field and recursive simulation of far-field tail flame interpolation, and can provide data support for research of infrared radiation characteristics, aerodynamic effects and the like of a target.
The technical scheme of the invention is as follows:
the hypersonic cruise aircraft full flow field numerical simulation method specifically comprises the following steps:
(1) Building a hypersonic cruise aircraft geometric model: based on the physical structure of the hypersonic cruise aircraft and the tail flame distribution rule, constructing an elastomer and near-field tail flame integrated geometric model and a far-field tail flame geometric model of the hypersonic cruise aircraft;
(2) Dividing grids: the projectile body and near-field tail flame integrated geometric model adopts unstructured grids to carry out grid division, and the far-field tail flame geometric model adopts structural grids to carry out grid division;
(3) Numerical simulation of the integration of the projectile body and the near-field tail flame flow field: firstly, carrying out numerical simulation calculation on grid division areas of an projectile and near-field tail flame integrated geometric model according to an ideal gas flow field, then carrying out numerical simulation calculation on a real gas component flow field according to a numerical simulation calculation result of the projectile and near-field tail flame ideal gas flow field and a real gas component, and finally carrying out numerical simulation calculation on a reburning reaction flow field according to a numerical simulation calculation result of the projectile and near-field tail flame real gas component flow field to obtain a numerical simulation calculation result of the projectile and near-field tail flame reburning reaction flow field, thereby completing integral numerical simulation of the projectile and the near-field tail flame flow field;
(4) Numerical simulation of far-field tail flame flow field: according to the numerical simulation calculation results of the projectile body and the near-field tail flame reburning reaction flow field outlet, an interpolation algorithm is adopted to obtain initialization parameters of the far-field tail flame flow field inlet, after numerical simulation calculation is carried out on the initialization parameters of the far-field tail flame flow field inlet according to the ideal gas flow field, numerical simulation calculation results of the far-field tail flame ideal gas flow field are obtained, then the interpolation algorithm is adopted to obtain the mass fraction of the real gas component at the far-field tail flame flow field inlet, then the numerical simulation calculation of the real gas component flow field is carried out according to the numerical simulation calculation results of the far-field tail flame ideal gas flow field and the mass fraction of the real gas component at the far-field tail flame flow field inlet, and finally the numerical simulation calculation of the reburning reaction flow field is carried out according to the numerical simulation calculation results of the far-field tail flame real gas component flow field, so that the numerical simulation calculation results of the far-field tail flame reburning reaction flow field are obtained.
In the step (2), the projectile body and near-field tail flame integrated geometric model is imported into grid division software, the projectile body and near-field tail flame integrated geometric model is divided into the regions near the projectile body surface and near the tail flame axis by using fine grids, and the regions far away from the projectile body surface and the tail flame axis are divided by using sparse grids, so that the sparse grids are coated on the periphery of the fine grids.
In the step (2), a far-field tail flame geometric model is led into grid division software, grids are gradually drawn in a sparse mode by using an equal-proportion algorithm along the tail flame axial direction from a region close to the spray pipe to a region far away from the spray pipe, grids are gradually drawn in a radial region vertical to the tail flame axial direction by using the equal-proportion algorithm from inside to outside, and after grid division is completed, the boundary of the far-field tail flame geometric model is further encrypted by using Fluent software.
The specific steps of the step (3) are as follows:
(a) The integrated geometric model of the projectile body and the near-field tail flame after grid division further encrypts the boundary positions of the projectile body wall surface, the engine spraying pipe wall surface and the sparse grid in the integrated geometric model by using Fluent software, then selects a density-based solver and sets related parameters and models;
(b) Setting the middle cross section of an engine spray pipe as a pressure inlet boundary in Fluent software, assigning a combustion material of the engine combustion chamber as ideal gas according to the combustion chamber parameters as pressure and temperature parameters of the pressure inlet boundary, setting the boundary of a sparse grid of an projectile and near-field tail flame integrated geometric model as a pressure far-field boundary, setting gas in the projectile and near-field tail flame integrated geometric model as ideal gas, and setting parameters and convergence conditions of an ideal gas flow field in the pressure far-field boundary; initializing a flow field according to parameters of an ideal gas flow field in a pressure far-field boundary, and then adopting a density-based solver to carry out iterative solution to obtain a numerical simulation calculation result of the projectile body and near-field tail flame ideal gas flow field;
(c) Setting a pressure inlet boundary condition of an engine combustion chamber according to the actual component composition and mass fraction of an engine combustion product, setting a pressure far-field boundary condition and a pressure outlet boundary condition according to the different gas component compositions and mass fraction in real gas, and taking the pressure inlet boundary condition, the pressure far-field boundary condition and the pressure outlet boundary condition as iteration conditions of the numerical simulation and calculation of the real gas component flow fields of the projectile body and the near-field tail flame; then setting convergence conditions of the projectile body and the near-field tail flame real gas component flow field, initializing the flow field according to iteration conditions of numerical simulation and analog computation of the projectile body and the near-field tail flame real gas component flow field, and carrying out iterative solution by adopting a density-based solver in combination with numerical simulation and analog computation results of the projectile body and the near-field tail flame ideal gas flow field to obtain numerical simulation and analog computation results of the projectile body and the near-field tail flame real gas component flow field;
(d) Setting a reburning reaction equation and condition parameters among all components in the tail flame as iteration conditions of numerical simulation calculation of the projectile and near-field tail flame reburning reaction flow field according to actual reaction conditions of all components in the tail flame, setting convergence conditions of the reburning reaction flow field, combining the iteration conditions of numerical simulation calculation of the projectile and near-field tail flame reburning reaction flow field with the numerical simulation calculation results of the projectile and near-field tail flame real gas component flow field, and adopting a density-based solver to carry out iterative solution to obtain the numerical simulation calculation results of the projectile and near-field tail flame reburning reaction flow field.
The specific steps of the step (4) are as follows:
(a) Firstly, a Fluent software reads parameter values of an elastomer and a near-field tail flame reburning reaction flow field outlet section, then calculates the parameter values of each node at a far-field tail flame inlet in an ideal gas flow field by using an interpolation algorithm, so as to assign a value to a far-field tail flame inlet boundary of the ideal gas flow field, simultaneously sets a far-field tail flame outlet boundary and a radial area boundary axially extending along the far-field tail flame, sets a gas environment of the far-field tail flame as ideal gas, sets the far-field tail flame outlet boundary and the radial area boundary as a pressure far-field boundary of the far-field tail flame ideal gas flow field, and then adopts a density-based solver to carry out iterative solution after initializing the flow field according to the pressure far-field boundary of the far-field tail flame ideal gas flow field, so as to obtain a numerical simulation calculation result of the far-field tail flame ideal gas flow field;
(b) Calculating the mass fraction of each node real gas component at the far-field tail flame inlet based on different gas component composition and mass fraction in the real gas of the near-field tail flame by using an interpolation algorithm, so as to assign a value to the far-field tail flame inlet boundary of the real gas component flow field, assign a value to the far-field tail flame outlet boundary of the real gas component flow field and a radial area boundary extending along the far-field tail flame axial direction according to the composition and mass fraction in the atmosphere, initialize the flow field according to the far-field tail flame outlet boundary of the real gas component flow field, and iteratively solve by adopting a density-based solver in combination with the numerical simulation calculation result of the far-field tail flame ideal gas flow field to obtain the numerical simulation calculation result of the far-field tail flame real gas component flow field;
(c) Setting a re-combustion reaction equation and condition parameters among all components in the tail flame as iteration conditions of numerical simulation and analog calculation of a far-field tail flame re-combustion reaction flow field according to actual reaction conditions of all the components in the tail flame, setting convergence conditions of the re-combustion reaction flow field, combining the iteration conditions of the numerical simulation and analog calculation of the far-field tail flame re-combustion reaction flow field with the numerical simulation and analog calculation results of a far-field tail flame real gas component flow field, and adopting a density-based solver to carry out iterative solution to obtain the numerical simulation and analog calculation results of the far-field tail flame re-combustion reaction flow field.
The interpolation algorithm is a cubic spline interpolation algorithm.
The hypersonic cruise aircraft full flow field numerical simulation system comprises an projectile and near-field tail flame integrated geometric model, a far-field tail flame geometric model, an unstructured grid division model, a structured grid division model and a full flow field numerical simulation model, wherein the full flow field numerical simulation model comprises an projectile and near-field tail flame flow field integrated numerical simulation model and a far-field tail flame flow field numerical simulation model; the projectile and near-field tail flame integrated geometric model adopts an unstructured grid division model to carry out unstructured grid division, and then the projectile and near-field tail flame integrated geometric model after grid division carries out numerical simulation operation through a projectile and near-field tail flame flow field integrated numerical simulation model to obtain a projectile and near-field tail flame flow field integrated numerical simulation result; the far-field tail flame geometric model adopts a structural meshing model to carry out structural meshing, and then the far-field tail flame geometric model after meshing carries out numerical simulation operation through a far-field tail flame flow field numerical simulation model to obtain a far-field tail flame flow field numerical simulation result.
The projectile and near-field tail flame flow field integrated numerical simulation model comprises a projectile and near-field tail flame ideal gas flow field integrated numerical simulation model, a projectile and near-field tail flame real gas component flow field integrated numerical simulation model and a projectile and near-field tail flame reburning reaction flow field integrated numerical simulation model.
The far-field tail flame flow field numerical simulation model comprises a far-field tail flame inlet interpolation algorithm model, a far-field tail flame ideal gas flow field numerical simulation model, a far-field tail flame real gas component flow field numerical simulation model and a far-field tail flame re-combustion reaction flow field numerical simulation model.
The invention has the advantages that:
(1) According to the hypersonic cruise aircraft, the unstructured grid and the structural grid are combined, and the hypersonic cruise aircraft body has a complex geometric structure, so that the body and near-field tail flame integrated geometric model adopts the unstructured grid to carry out grid division, the far-field tail flame geometric model is relatively simple, and the far-field tail flame geometric model adopts the structural grid to carry out grid division, so that the grid division difficulty, the workload and the numerical simulation efficiency are both considered.
(2) According to the invention, the coupling effect of the projectile body and the near-field tail flame flow field can be accurately simulated by carrying out integrated numerical simulation on the projectile body and the near-field tail flame flow field, so that the distribution situation of the projectile body and the near-field tail flame flow field can be accurately obtained, and simulation results show that the projectile body has a significant influence on the distribution of the tail flame flow field, if the coupling effect is not considered, a large error occurs in the calculation result of the tail flame flow field, and after the coupling effect of the projectile body and the near-field tail flame is considered, the tail flame temperature is significantly increased and the width is significantly increased compared with the case of not considering the coupling effect.
(3) The method comprises the steps of carrying out integral numerical simulation on an elastomer and a near-field tail flame and carrying out numerical simulation on a far-field tail flame flow field by adopting gradual initialization, namely, firstly, assuming gas as ideal gas to obtain a numerical simulation calculation result of an ideal gas flow field, then taking the result as an initialization condition of simulation of a real gas component flow field, carrying out numerical simulation calculation on mass fractions of the real gas component flow field to obtain a numerical simulation calculation result of the real gas component flow field after initializing, carrying out numerical simulation calculation by combining a tail flame re-combustion reaction condition, and finally obtaining the numerical simulation calculation result of a re-combustion reaction flow field, so that the number of iterations of each numerical simulation calculation is small, the simulation calculation is easy to converge, and the accuracy of the numerical simulation calculation result is high.
(4) When the far-field tail flame flow field value simulation is carried out, the heterogeneous grid recursion assignment calculation method based on the interpolation algorithm is adopted, so that when the near-field tail flame grid accuracy is different from the far-field tail flame grid accuracy, the grid data are spliced, and the efficiency is improved on the premise of ensuring the simulation result accuracy.
Drawings
FIG. 1 is a block diagram of a hypersonic cruise vehicle full flow field numerical simulation system.
FIG. 2 is a schematic diagram of a grid-partitioned structure in a geometry model of an elastomer and near-field tail flame integration of the present invention.
FIG. 3 is a graph of a comparison of simulated temperature of an elastomer and near-field tail flow field integrated numerical values at a distance of 12m from the outlet of the tail flame nozzle with simulated temperature of an individual tail flow field in an embodiment of the present invention.
Detailed Description
The following description of the embodiments of the present invention will be made clearly and completely with reference to the accompanying drawings, in which it is apparent that the embodiments described are only some embodiments of the present invention, but not all embodiments. All other embodiments, which can be made by those skilled in the art based on the embodiments of the invention without making any inventive effort, are intended to be within the scope of the invention.
Referring to fig. 1, a hypersonic cruise aircraft full flow field numerical simulation system comprises an projectile and near-field tail flame integrated geometric model 1, a far-field tail flame geometric model 2, an unstructured grid division model 3, a structured grid division model 4 and a full flow field numerical simulation model 5, wherein the full flow field numerical simulation model 5 comprises a projectile and near-field tail flame flow field integrated numerical simulation model 51 and a far-field tail flame flow field numerical simulation model 52; the projectile and near-field tail flame integrated geometric model 1 adopts an unstructured grid division model 3 to carry out unstructured grid division, and then the projectile and near-field tail flame integrated geometric model after grid division carries out numerical simulation operation through a projectile and near-field tail flame flow field integrated numerical simulation model 51 to obtain a projectile and near-field tail flame flow field integrated numerical simulation result; the far-field tail flame geometric model 2 adopts the structural meshing model 4 to carry out structural meshing, and then the far-field tail flame geometric model after meshing carries out numerical simulation operation through the far-field tail flame flow field numerical simulation model 52 to obtain a far-field tail flame flow field numerical simulation result.
The hypersonic cruise aircraft full flow field numerical simulation method specifically comprises the following steps:
(1) Building a hypersonic cruise aircraft geometric model: based on the physical structure of the hypersonic cruise aircraft and the tail flame distribution rule, constructing an elastomer and near-field tail flame integrated geometric model and a far-field tail flame geometric model of the hypersonic cruise aircraft;
(2) Dividing grids: referring to fig. 2, a geometry model of an elastomer and a near-field tail flame integration is imported into Mesh software, a cylindrical region 12 near the surface of an elastomer 11 and near the axis of the tail flame is divided by using a fine grid, a cylindrical region 13 far from the surface of the elastomer and the axis of the tail flame is divided by using a sparse grid, and the periphery of the fine grid is covered by the sparse grid; introducing far-field tail flame geometric model 2 into ICEM software, gradually sparsely drawing grids along the tail flame axial direction from a region close to the spray pipe to a region far from the spray pipe by using an equal-proportion algorithm, gradually sparsely drawing grids along the radial region vertical to the tail flame axial direction from inside to outside by using the equal-proportion algorithm, and further encrypting the boundary of far-field tail flame geometric model 2 by using Fluent software after grid division is completed;
(3) The numerical simulation of the integration of the projectile body and the near-field tail flame flow field comprises the following steps:
(a) The integrated geometry model 1 of the projectile body and the near-field tail flame after grid division further encrypts the boundary positions of the projectile body wall surface, the engine spraying pipe wall surface and the sparse grid in the integrated geometry model 1 by using Fluent software, then selects a density-based solver and sets related parameters and models (including an implicit solving mode, a standard k-epsilon equation turbulence model, a laminar flow finite transportation model and the like);
(b) Setting the middle cross section of an engine spray pipe as a pressure inlet boundary in Fluent software, assigning a combustion material of the engine combustion chamber as ideal gas according to the combustion chamber parameters as pressure and temperature parameters of the pressure inlet boundary, setting the boundary of a sparse grid of an elastomer and near-field tail flame integrated geometric model as a pressure far-field boundary, setting gas in the elastomer and near-field tail flame integrated geometric model as ideal gas, and setting parameters (pressure, temperature, speed and the like of ideal gas) and convergence conditions of an ideal gas flow field in the pressure far-field boundary;
the flow field of the X-43A aircraft under Mach 6 is simulated, and simulation parameters are as follows:
TABLE 1 spray tube inlet set-up under 6Ma conditions
| Variable name | Parameter value | Variable name | Parameter value |
| Total pressure of nozzle inlet | 386729Pa | Static pressure at the nozzle inlet | 98717Pa |
| Total temperature of nozzle inlet | 2218K | Ambient static temperature | 223K |
| Static pressure of environment | 1900Pa | Fly height | 26944m |
Initializing a flow field according to parameters of an ideal gas flow field in a pressure far-field boundary, and then adopting a density-based solver to carry out iterative solution to obtain numerical simulation calculation results (including temperature, pressure, speed, turbulence parameters and the like) of the projectile body and near-field tail flame ideal gas flow field;
(c) Setting a pressure inlet boundary condition of an engine combustion chamber according to the actual component composition and mass fraction of an engine combustion product, setting a pressure far-field boundary condition and a pressure outlet boundary condition according to the different gas component compositions and mass fraction in real gas, and taking the pressure inlet boundary condition, the pressure far-field boundary condition and the pressure outlet boundary condition as iteration conditions of the numerical simulation and calculation of the real gas component flow fields of the projectile body and the near-field tail flame;
TABLE 2 mass fractions of the components of the real gas at the nozzle inlet and in the atmosphere under 6Ma conditions
Then setting convergence conditions of the projectile body and the near-field tail flame real gas component flow field, initializing the flow field according to iteration conditions of numerical simulation and analog computation of the projectile body and the near-field tail flame real gas component flow field, and carrying out iterative solution by adopting a density-based solver in combination with numerical simulation and analog computation results of the projectile body and the near-field tail flame ideal gas flow field to obtain numerical simulation and analog computation results of the projectile body and the near-field tail flame real gas component flow field;
as shown in fig. 3, the dashed line is the temperature distribution of the integral simulation of the projectile body and the near-field tail flame flow field, the solid line is the temperature distribution of the independent simulation tail flame flow field, and the coupling effect is ignored, as can be found from fig. 3, when the coupling effect is ignored, the tail flame calculation result has larger error, in particular, the temperature is lower and the size is smaller (the highest temperature of the tail flame of the integral simulation of the projectile body and the near-field tail flame flow field is 526.3, the highest temperature of the tail flame of the independent simulation is 441.1, which is improved by 19% compared with the highest temperature of the tail flame, and the width of the tail flame is obviously widened);
(d) Setting a re-combustion reaction equation and condition parameters (such as a front factor, an activation energy, a temperature fraction and the like) among all components in the tail flame as iteration conditions of numerical simulation calculation of an projectile body and a near-field tail flame re-combustion reaction flow field according to actual reaction conditions of all the components in the tail flame, setting convergence conditions of the re-combustion reaction flow field, combining the iteration conditions of the numerical simulation calculation of the projectile body and the near-field tail flame re-combustion reaction flow field with the numerical simulation calculation results of the projectile body and the near-field tail flame real gas component flow field, and adopting a density-based solver to carry out iterative solution to obtain the numerical simulation calculation results of the projectile body and the near-field tail flame re-combustion reaction flow field;
(4) The structured grids of the far-field tail flame geometric model are only 90 grids at the inlet section, which is far less than 340 grids of the outlet section of the projectile body and near-field tail flame integrated geometric model, and the corresponding relation among the grids also has larger gap, so that larger problems can be generated in the initial stage of simulation; therefore, by analyzing the numerical simulation principle, grid calculation is essentially discretization of a physical continuous field, numerical simulation is carried out on different grid divisions, only represents approximate simplified divisions of the original continuous distribution physical field, therefore, no rigidification is required to be completely consistent before and after, only reasonable values are required to be assigned to far-field tail flame inlet grids according to data at the near-field tail flame grid outlet, and according to the principle, the establishment of a corresponding relation between two irrelevant grids is completely feasible. Therefore, a cubic spline interpolation algorithm is adopted to try to establish the corresponding relation between grids with completely different radial node parameters, so that the calculation efficiency is improved, and the high calculation precision is ensured;
the numerical simulation of the far-field tail flame flow field specifically comprises the following steps:
(a) Firstly, a Fluent software reads parameter values of an elastomer and a near-field tail flame reburning reaction flow field outlet section, a UDF program is utilized to write a cubic spline interpolation algorithm, then the parameter values of the elastomer and the near-field tail flame reburning reaction flow field outlet section are calculated by using the cubic spline interpolation algorithm to obtain parameter values of each node at an ideal gas flow field at a far-field tail flame inlet, so that an ideal gas flow field far-field tail flame inlet boundary is assigned, a far-field tail flame outlet boundary and a radial area boundary axially extending along the far-field tail flame are set, the gas environment of the far-field tail flame is set as ideal gas, the far-field tail flame outlet boundary and the radial area boundary are set as pressure far-field boundaries of the far-field tail flame ideal gas flow field, and then a density-based solver is adopted to carry out iterative solution after initializing the flow field according to the pressure far-field boundary of the far-field tail flame ideal gas flow field, so that a numerical simulation calculation result of the far-field tail flame ideal gas flow field is obtained;
(b) Calculating the mass fraction of each node real gas component at the far-field tail flame inlet based on different gas component composition and mass fraction in the real gas of the near-field tail flame by using an interpolation algorithm, so as to assign a value to the far-field tail flame inlet boundary of the real gas component flow field, assign a value to the far-field tail flame outlet boundary of the real gas component flow field and a radial area boundary extending along the far-field tail flame axial direction according to the composition and mass fraction in the atmosphere, initialize the flow field according to the far-field tail flame outlet boundary of the real gas component flow field, and iteratively solve by adopting a density-based solver in combination with the numerical simulation calculation result of the far-field tail flame ideal gas flow field to obtain the numerical simulation calculation result of the far-field tail flame real gas component flow field;
(c) Setting a re-combustion reaction equation and condition parameters among all components in the tail flame as iteration conditions of numerical simulation and analog calculation of a far-field tail flame re-combustion reaction flow field according to actual reaction conditions of all the components in the tail flame, setting convergence conditions of the re-combustion reaction flow field, combining the iteration conditions of the numerical simulation and analog calculation of the far-field tail flame re-combustion reaction flow field with the numerical simulation and analog calculation results of a far-field tail flame real gas component flow field, and adopting a density-based solver to carry out iterative solution to obtain the numerical simulation and analog calculation results of the far-field tail flame re-combustion reaction flow field.
Although embodiments of the present invention have been shown and described, it will be understood by those skilled in the art that various changes, modifications, substitutions and alterations can be made therein without departing from the principles and spirit of the invention, the scope of which is defined in the appended claims and their equivalents.
Claims (9)
1. A hypersonic cruise aircraft full flow field numerical simulation method is characterized in that: the method specifically comprises the following steps:
(1) Building a hypersonic cruise aircraft geometric model: based on the physical structure of the hypersonic cruise aircraft and the tail flame distribution rule, constructing an elastomer and near-field tail flame integrated geometric model and a far-field tail flame geometric model of the hypersonic cruise aircraft;
(2) Dividing grids: the projectile body and near-field tail flame integrated geometric model adopts unstructured grids to carry out grid division, and the far-field tail flame geometric model adopts structural grids to carry out grid division;
(3) Numerical simulation of the integration of the projectile body and the near-field tail flame flow field: firstly, carrying out numerical simulation calculation on grid division areas of an projectile and near-field tail flame integrated geometric model according to an ideal gas flow field, then carrying out numerical simulation calculation on a real gas component flow field according to a numerical simulation calculation result of the projectile and near-field tail flame ideal gas flow field and a real gas component, and finally carrying out numerical simulation calculation on a reburning reaction flow field according to a numerical simulation calculation result of the projectile and near-field tail flame real gas component flow field to obtain a numerical simulation calculation result of the projectile and near-field tail flame reburning reaction flow field, thereby completing integral numerical simulation of the projectile and the near-field tail flame flow field;
(4) Numerical simulation of far-field tail flame flow field: according to the numerical simulation calculation results of the projectile body and the near-field tail flame reburning reaction flow field outlet, an interpolation algorithm is adopted to obtain initialization parameters of the far-field tail flame flow field inlet, after numerical simulation calculation is carried out on the initialization parameters of the far-field tail flame flow field inlet according to the ideal gas flow field, numerical simulation calculation results of the far-field tail flame ideal gas flow field are obtained, then the interpolation algorithm is adopted to obtain the mass fraction of the real gas component at the far-field tail flame flow field inlet, then the numerical simulation calculation of the real gas component flow field is carried out according to the numerical simulation calculation results of the far-field tail flame ideal gas flow field and the mass fraction of the real gas component at the far-field tail flame flow field inlet, and finally the numerical simulation calculation of the reburning reaction flow field is carried out according to the numerical simulation calculation results of the far-field tail flame real gas component flow field, so that the numerical simulation calculation results of the far-field tail flame reburning reaction flow field are obtained.
2. The hypersonic cruise aircraft full flow field numerical simulation method according to claim 1, wherein the method is characterized in that: in the step (2), the projectile body and near-field tail flame integrated geometric model is imported into grid division software, the projectile body and near-field tail flame integrated geometric model is divided into the regions near the projectile body surface and near the tail flame axis by using fine grids, and the regions far away from the projectile body surface and the tail flame axis are divided by using sparse grids, so that the sparse grids are coated on the periphery of the fine grids.
3. The hypersonic cruise aircraft full flow field numerical simulation method according to claim 1, wherein the method is characterized in that: in the step (2), a far-field tail flame geometric model is led into grid division software, grids are gradually drawn in a sparse mode by using an equal-proportion algorithm along the tail flame axial direction from a region close to the spray pipe to a region far away from the spray pipe, grids are gradually drawn in a radial region vertical to the tail flame axial direction by using the equal-proportion algorithm from inside to outside, and after grid division is completed, the boundary of the far-field tail flame geometric model is further encrypted by using Fluent software.
4. The hypersonic cruise aircraft full flow field numerical simulation method according to claim 1, wherein the method is characterized in that: the specific steps of the step (3) are as follows:
(a) The integrated geometric model of the projectile body and the near-field tail flame after grid division further encrypts the boundary positions of the projectile body wall surface, the engine spraying pipe wall surface and the sparse grid in the integrated geometric model by using Fluent software, then selects a density-based solver and sets related parameters and models;
(b) Setting the middle cross section of an engine spray pipe as a pressure inlet boundary in Fluent software, assigning a combustion material of the engine combustion chamber as ideal gas according to the combustion chamber parameters as pressure and temperature parameters of the pressure inlet boundary, setting the boundary of a sparse grid of an projectile and near-field tail flame integrated geometric model as a pressure far-field boundary, setting gas in the projectile and near-field tail flame integrated geometric model as ideal gas, and setting parameters and convergence conditions of an ideal gas flow field in the pressure far-field boundary; initializing a flow field according to parameters of an ideal gas flow field in a pressure far-field boundary, and then adopting a density-based solver to carry out iterative solution to obtain a numerical simulation calculation result of the projectile body and near-field tail flame ideal gas flow field;
(c) Setting a pressure inlet boundary condition of an engine combustion chamber according to the actual component composition and mass fraction of an engine combustion product, setting a pressure far-field boundary condition and a pressure outlet boundary condition according to the different gas component compositions and mass fraction in real gas, and taking the pressure inlet boundary condition, the pressure far-field boundary condition and the pressure outlet boundary condition as iteration conditions of the numerical simulation and calculation of the real gas component flow fields of the projectile body and the near-field tail flame; then setting convergence conditions of the projectile body and the near-field tail flame real gas component flow field, initializing the flow field according to iteration conditions of numerical simulation and analog computation of the projectile body and the near-field tail flame real gas component flow field, and carrying out iterative solution by adopting a density-based solver in combination with numerical simulation and analog computation results of the projectile body and the near-field tail flame ideal gas flow field to obtain numerical simulation and analog computation results of the projectile body and the near-field tail flame real gas component flow field;
(d) Setting a reburning reaction equation and condition parameters among all components in the tail flame as iteration conditions of numerical simulation calculation of the projectile and near-field tail flame reburning reaction flow field according to actual reaction conditions of all components in the tail flame, setting convergence conditions of the reburning reaction flow field, combining the iteration conditions of numerical simulation calculation of the projectile and near-field tail flame reburning reaction flow field with the numerical simulation calculation results of the projectile and near-field tail flame real gas component flow field, and adopting a density-based solver to carry out iterative solution to obtain the numerical simulation calculation results of the projectile and near-field tail flame reburning reaction flow field.
5. The hypersonic cruise aircraft full flow field numerical simulation method according to claim 1, wherein the method is characterized in that: the specific steps of the step (4) are as follows:
(a) Firstly, a Fluent software reads parameter values of an elastomer and a near-field tail flame reburning reaction flow field outlet section, then calculates the parameter values of each node at a far-field tail flame inlet in an ideal gas flow field by using an interpolation algorithm, so as to assign a value to a far-field tail flame inlet boundary of the ideal gas flow field, simultaneously sets a far-field tail flame outlet boundary and a radial area boundary axially extending along the far-field tail flame, sets a gas environment of the far-field tail flame as ideal gas, sets the far-field tail flame outlet boundary and the radial area boundary as a pressure far-field boundary of the far-field tail flame ideal gas flow field, and then adopts a density-based solver to carry out iterative solution after initializing the flow field according to the pressure far-field boundary of the far-field tail flame ideal gas flow field, so as to obtain a numerical simulation calculation result of the far-field tail flame ideal gas flow field;
(b) Calculating the mass fraction of each node real gas component at the far-field tail flame inlet based on different gas component composition and mass fraction in the real gas of the near-field tail flame by using an interpolation algorithm, so as to assign a value to the far-field tail flame inlet boundary of the real gas component flow field, assign a value to the far-field tail flame outlet boundary of the real gas component flow field and a radial area boundary extending along the far-field tail flame axial direction according to the composition and mass fraction in the atmosphere, initialize the flow field according to the far-field tail flame outlet boundary of the real gas component flow field, and iteratively solve by adopting a density-based solver in combination with the numerical simulation calculation result of the far-field tail flame ideal gas flow field to obtain the numerical simulation calculation result of the far-field tail flame real gas component flow field;
(c) Setting a re-combustion reaction equation and condition parameters among all components in the tail flame as iteration conditions of numerical simulation and analog calculation of a far-field tail flame re-combustion reaction flow field according to actual reaction conditions of all the components in the tail flame, setting convergence conditions of the re-combustion reaction flow field, combining the iteration conditions of the numerical simulation and analog calculation of the far-field tail flame re-combustion reaction flow field with the numerical simulation and analog calculation results of a far-field tail flame real gas component flow field, and adopting a density-based solver to carry out iterative solution to obtain the numerical simulation and analog calculation results of the far-field tail flame re-combustion reaction flow field.
6. The hypersonic cruise aircraft full flow field numerical simulation method according to claim 1 or 5, wherein the method comprises the following steps of: the interpolation algorithm is a cubic spline interpolation algorithm.
7. A hypersonic cruise aircraft full flow field numerical simulation system is characterized in that: the system comprises a projectile body and near-field tail flame integrated geometric model, a far-field tail flame geometric model, an unstructured grid division model, a structural grid division model and a full flow field numerical simulation model, wherein the full flow field numerical simulation model comprises a projectile body and near-field tail flame flow field integrated numerical simulation model and a far-field tail flame flow field numerical simulation model;
the method comprises the steps that the non-structural grid division model is adopted by the projectile and near-field tail flame integrated geometric model to carry out non-structural grid division, the grid division area of the projectile and near-field tail flame integrated geometric model is firstly subjected to numerical simulation and analog calculation according to an ideal gas flow field, then the numerical simulation and analog calculation of a real gas component flow field is carried out according to the numerical simulation and analog calculation result of the projectile and near-field tail flame ideal gas flow field and the real gas component, finally the numerical simulation and analog calculation of a re-combustion reaction flow field is carried out according to the numerical simulation and analog calculation result of the projectile and near-field tail flame re-combustion reaction flow field, and the numerical simulation and analog calculation result of the projectile and near-field tail flame flow field integrated numerical simulation is obtained;
the far-field tail flame geometric model adopts a structural grid division model to carry out structural grid division, the far-field tail flame flow field numerical simulation model obtains initialization parameters at the far-field tail flame flow field inlet according to numerical simulation calculation results of an projectile body and a near-field tail flame afterburning reaction flow field outlet by adopting an interpolation algorithm, the initialization parameters at the far-field tail flame flow field inlet are subjected to numerical simulation calculation according to an ideal gas flow field to obtain numerical simulation calculation results of a far-field tail flame ideal gas flow field, then real gas components at the projectile body and the near-field tail flame afterburning reaction flow field outlet are subjected to interpolation algorithm to obtain mass fractions of real gas components at the far-field tail flame flow field inlet, numerical simulation calculation of a real gas component flow field is carried out according to the numerical simulation calculation results of the far-field tail flame ideal gas components flow field, and finally the numerical simulation calculation results of the afterburning reaction flow field are carried out according to the numerical simulation calculation results of the far-field tail flame real gas components flow field, and the numerical simulation calculation results of the far-field afterburning reaction flow field is obtained.
8. The hypersonic cruise aircraft full flow field numerical simulation system as set forth in claim 7, wherein: the projectile and near-field tail flame flow field integrated numerical simulation model comprises a projectile and near-field tail flame ideal gas flow field integrated numerical simulation model, a projectile and near-field tail flame real gas component flow field integrated numerical simulation model and a projectile and near-field tail flame reburning reaction flow field integrated numerical simulation model.
9. The hypersonic cruise aircraft full flow field numerical simulation system as set forth in claim 7, wherein: the far-field tail flame flow field numerical simulation model comprises a far-field tail flame inlet interpolation algorithm model, a far-field tail flame ideal gas flow field numerical simulation model, a far-field tail flame real gas component flow field numerical simulation model and a far-field tail flame re-combustion reaction flow field numerical simulation model.
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