A coupled simulation method, apparatus, equipment, and medium for multiple components and multiple solvers of aero-engines.

Abstract

This invention relates to the field of aero-engine technology, and discloses a method, apparatus, equipment, and medium for coupled simulation of multiple components and multiple solvers in aero-engines. This invention precisely matches the RANS steady turbulence model and the LES large eddy simulation solver to the flow field characteristics of different components such as fans / compressors, combustion chambers, turbines, afterburners, and nozzles. It solves the problem that traditional single models or solvers cannot adapt to multi-physics coupling, thus improving the simulation accuracy of the entire aero-engine. This invention adopts a strategy of "component-level decomposition + parallel multi-solver + model adaptation," significantly reducing the computational load caused by uniformly using high-precision LES simulation throughout the entire process. Simultaneously, it ensures the coupling accuracy between components through interface data collaborative processing, improving the simulation efficiency of the entire aero-engine and achieving a balance between simulation efficiency and accuracy. This effectively solves the problem of dynamic coupling of flow fields between aero-engine components, enhancing the engineering applicability of this invention.

Description

Technical Field

[0001] This invention relates to the field of aero-engine simulation technology, and discloses a method, apparatus, equipment, and medium for coupled simulation of multiple components and multiple solvers of aero-engines. Background Technology

[0002] The internal flow field of an aero-engine is a typical multiphysics coupled system, encompassing complex phenomena such as turbulence, combustion, phase change, shock waves, and heat conduction. Furthermore, strong aerodynamic and thermodynamic coupling exists between various components (such as the fan and compressor, and the combustion chamber and turbine). The flow field parameters of upstream components (such as pressure and temperature) directly determine the working boundaries of downstream components, and the flow field feedback of downstream components (such as backflow and pressure loss) also affects the stable operation of upstream components. Therefore, full-process, high-precision coupled simulation is a core technological support for engine development.

[0003] However, current aero-engine simulation methods, some of which use RANS (Reynolds-averaged Navier-Stokes) solvers or single turbulence models for each engine component, while having high computational efficiency, lack sufficient accuracy in simulating complex flow fields of some components, and cannot support the refined design of key components, resulting in a serious imbalance between simulation efficiency and accuracy. Summary of the Invention

[0004] The purpose of this invention is to provide a multi-component, multi-solver coupled simulation method, device, equipment, and medium for aero-engines, which matches corresponding turbulence models and solvers for different components to improve the accuracy and efficiency of the whole-engine simulation, and realizes dynamic coupling of flow fields between components through the interface data processing mechanism, thereby improving the consistency between simulation results and real physical processes.

[0005] To achieve the above-mentioned technical effects, the technical solution adopted by the present invention is as follows:

[0006] A coupled simulation method for multiple components and multiple solvers of an aero-engine includes:

[0007] Simulation models of each component of the aero-engine are obtained and meshed, and the boundary conditions for the simulation of the entire aero-engine are determined. The components of the aero-engine include the fan, compressor, bypass duct, main combustion chamber, high-pressure turbine, low-pressure turbine, afterburner, and nozzle. The boundary conditions include the total temperature and pressure at the inlet of the aero-engine, the incoming Mach number, and the back pressure at the nozzle outlet.

[0008] Based on the flow field characteristics of each component of the aero-engine, the corresponding turbulence model and solver are matched for the simulation model of each component;

[0009] Based on the boundary conditions of the aero-engine whole-machine simulation, a radial-circumferential mesh mapping algorithm is used to transfer the flow field parameters of the grid points on the interface outlet plane of the upstream component to the grid points on the interface inlet plane of the downstream component. The solvers of each component are driven sequentially according to the airflow sequence of the aero-engine to simulate and calculate the flow field in each component of the aero-engine, and obtain the whole-machine simulation results of the aero-engine. The flow field parameters include radial velocity, circumferential velocity, axial velocity, temperature, pressure, and component concentration. The simulation results include the thrust value of the aero-engine, the efficiency of each component, and the combustion chamber temperature distribution.

[0010] Furthermore, the solvers of each component are driven sequentially according to the airflow sequence of the aero-engine. When simulating the flow field within each component, the flow field parameters between the inlet and outlet of each component are updated synchronously and iteratively until the inlet and outlet flow deviation of each component is less than the preset flow deviation threshold or the total temperature and total pressure deviation of the key inlet and outlet sections is less than the preset deviation threshold. Then, it is determined that the flow field iteration of the entire aero-engine has reached the convergence state, and the simulation results of the entire aero-engine are output.

[0011] Furthermore, the expression for transferring the flow field parameters of the upstream component at the grid points on the interface outlet plane to the downstream component at the grid points on the interface inlet plane is as follows:

[0012] ;

[0013] in: The coordinates of the grid points of the downstream component on the inlet plane of the interface; This is an empirical constant, with a value range of 0 to 1; The radial gradient of the flow field parameters at grid points on the inlet plane of the interface of the upstream component; These are the coordinates of the grid points of the upstream component on the inlet plane of the interface. , , , The upstream component grid points enclose the downstream grid points respectively. The coordinates of the four neighboring grid points; , , , , Grid points , , , , The flow field parameters.

[0014] Furthermore, the fan, compressor, and bypass duct adopt the SA turbulence model and the RANS steady solver based on the SA turbulence model; the main combustion chamber adopts the LES large eddy simulation unsteady solver; the high-pressure turbine, the low-pressure turbine, and the nozzle adopt the SST turbulence model and the RANS steady solver based on the SST turbulence model; and the afterburner adopts the LES large eddy simulation solver.

[0015] An aero-engine whole-machine simulation device, used to implement the aforementioned coupled simulation method for multiple components and multiple solvers of an aero-engine, includes:

[0016] The component-level simulation model acquisition module is used to acquire and mesh simulation models of various components of the aero-engine, and simultaneously determine the boundary conditions for the overall simulation of the aero-engine. The components of the aero-engine include the fan, compressor, bypass duct, main combustion chamber, high-pressure turbine, low-pressure turbine, afterburner, and nozzle. The boundary conditions include the total temperature and pressure at the inlet of the aero-engine, the incoming Mach number, and the back pressure at the nozzle outlet.

[0017] The turbulence model and solver configuration module is used to match the corresponding turbulence model and solver to the simulation model of each component based on the flow field characteristics of each component of the aero-engine.

[0018] The simulation execution module is used to transfer the flow field parameters of the grid points on the interface outlet plane of the upstream component to the grid points on the interface inlet plane of the downstream component based on the boundary conditions of the aero-engine whole-engine simulation using a radial-circumferential mesh mapping algorithm. It sequentially drives the solvers of each component according to the airflow sequence of the aero-engine to perform simulation calculations on the flow field within each component, obtaining the whole-engine simulation results. The flow field parameters include radial velocity, circumferential velocity, axial velocity, temperature, pressure, and component concentration. The simulation results include the thrust value of the aero-engine, the efficiency of each component, and the combustion chamber temperature distribution.

[0019] Furthermore, the solvers of each component are driven sequentially according to the airflow sequence of the aero-engine. When simulating the flow field within each component, the flow field parameters between the inlet and outlet of each component are updated synchronously and iteratively until the inlet and outlet flow deviation of each component is less than the preset flow deviation threshold or the total temperature and total pressure deviation of the key inlet and outlet sections is less than the preset deviation threshold. Then, it is determined that the flow field iteration of the entire aero-engine has reached the convergence state, and the simulation results of the entire aero-engine are output.

[0020] A computer device includes a memory, a processor, and a computer program stored in the memory and executable on the processor. When the processor executes the computer program, it implements the coupled simulation method for multiple components and multiple solvers of an aero-engine as described in any of the preceding claims.

[0021] A computer-readable storage medium stores a computer program that executes a coupled simulation method for multiple components and multiple solvers of an aero-engine as described in any of the preceding claims.

[0022] Compared with the prior art, the beneficial effects of this invention are:

[0023] This invention addresses the flow field characteristics of different components such as fans / compressors, combustion chambers, turbines, afterburners, and nozzles. It precisely matches the RANS steady turbulence model and the LES large eddy simulation solver, fully leveraging the advantages of each model to solve the problem that traditional single turbulence models or solvers cannot adapt to multi-physics coupling, thus improving the simulation accuracy of the entire aero-engine. This invention employs a strategy of "component-level decomposition + parallel multi-solver + model adaptation," significantly reducing the computational load caused by uniformly using high-precision LES simulation throughout the entire process and the simulation errors caused by a single turbulence model. Simultaneously, it ensures the coupling accuracy between components through a collaborative interface data processing method, improving the simulation efficiency of the entire aero-engine and achieving a balance between simulation efficiency and accuracy. This invention uses an interface radial-circumferential processing algorithm as the flow field parameter transfer mechanism, effectively solving the problem of dynamic flow field coupling between aero-engine components and enhancing the engineering applicability of this invention. Attached Figure Description

[0024] Figure 1 The flowchart is shown below for the coupled simulation method of multiple components and multiple solvers of the aero-engine in this embodiment.

[0025] Figure 2 This is a flowchart of the coupled simulation method for multiple components and multiple solvers of an aero-engine in the embodiment;

[0026] Figure 3 This is a structural block diagram of a multi-component, multi-solver coupled simulation device for an aero-engine, as shown in the embodiment.

[0027] Figure 4 This is a structural block diagram of the computer device in the embodiment;

[0028] Wherein, 1-memory, 2-processor. Detailed Implementation

[0029] The present invention will now be described in further detail with reference to the embodiments and accompanying drawings. However, this should not be construed as limiting the scope of the above-described subject matter of the present invention to the following embodiments; all technologies implemented based on the content of the present invention fall within the scope of the present invention.

[0030] Example 1

[0031] See Figure 1This embodiment provides a coupled simulation method for multiple components and multiple solvers of an aero-engine, including:

[0032] Step 1: Decompose the entire aero-engine into components such as fan, compressor, bypass duct, main combustion chamber, high-pressure turbine, low-pressure turbine, afterburner, and nozzle; obtain simulation models of each component of the aero-engine and mesh them, while determining the boundary conditions for the simulation of the entire aero-engine; the boundary conditions include parameters such as the total temperature and pressure at the inlet of the aero-engine, the incoming Mach number, and the nozzle exit back pressure.

[0033] Step 2: Based on the flow field characteristics of each component of the aero-engine, match the corresponding turbulence model and solver to the simulation model of each component;

[0034] Step 3: Based on the boundary conditions of the aero-engine whole-machine simulation, a radial-circumferential mesh mapping algorithm is used to transfer the flow field parameters of the grid points on the interface outlet plane of the upstream component to the grid points on the interface inlet plane of the downstream component. The solvers of each component are driven sequentially according to the airflow sequence of the aero-engine to simulate and calculate the flow field within each component. The flow field parameters between the inlet and outlet of each component are updated synchronously and iteratively until the inlet and outlet flow deviation of each component is less than the preset flow deviation threshold or the total temperature and total pressure deviation of the key inlet and outlet sections is less than the preset deviation threshold. Then, it is determined that the aero-engine whole-machine flow field iteration has reached the convergence state, and the whole-machine simulation results of the aero-engine are output. The flow field parameters include radial velocity, circumferential velocity, axial velocity, temperature, pressure, and component concentration. The simulation results include the thrust value of the aero-engine, the efficiency of each component, and the combustion chamber temperature distribution.

[0035] This invention addresses the flow field characteristics of different components such as fans / compressors, combustion chambers, turbines, afterburners, and nozzles. It precisely matches the RANS steady turbulence model and the LES large eddy simulation solver, fully leveraging the advantages of each model to solve the problem that traditional single turbulence models or solvers cannot handle multi-physics coupling, thus improving the simulation accuracy of the entire aero-engine. This invention employs a strategy of "component-level decomposition + parallel multi-solver + model adaptation," reducing unnecessary high-precision simulation calculations throughout the entire process. Simultaneously, it ensures the coupling accuracy between components through collaborative processing of interface data, improving the simulation efficiency of the entire aero-engine and achieving a balance between simulation efficiency and accuracy. This invention utilizes an interface radial-circumferential processing algorithm flow field parameter transfer mechanism, effectively solving the problem of dynamic flow field coupling between aero-engine components and enhancing the engineering applicability of this invention.

[0036] Based on the same inventive concept, see Figure 3 This embodiment also provides a coupled simulation device for multiple components and multiple solvers of an aero-engine, used to implement the aforementioned aero-engine whole-engine simulation method, including:

[0037] The component-level simulation model acquisition module is used to acquire and mesh simulation models of various components of the aero-engine, and simultaneously determine the boundary conditions for the overall simulation of the aero-engine. The components of the aero-engine include the fan, compressor, bypass duct, main combustion chamber, high-pressure turbine, low-pressure turbine, afterburner, and nozzle. The boundary conditions include parameters such as the total temperature and pressure at the inlet of the aero-engine, the incoming Mach number, and the back pressure at the nozzle exit.

[0038] The turbulence model and solver configuration module is used to match the corresponding turbulence model and solver to the simulation model of each component based on the flow field characteristics of each component of the aero-engine.

[0039] The simulation execution module is used to transfer the flow field parameters of the grid points on the interface outlet plane of the upstream component to the grid points on the interface inlet plane of the downstream component based on the boundary conditions of the aero-engine whole-engine simulation using a radial-circumferential mesh mapping algorithm. It sequentially drives the solvers of each component according to the airflow sequence of the aero-engine to perform simulation calculations on the flow field within each component, obtaining the whole-engine simulation results. The flow field parameters include radial velocity, circumferential velocity, axial velocity, temperature, pressure, and component concentration. The simulation results include the thrust value of the aero-engine, the efficiency of each component, and the combustion chamber temperature distribution.

[0040] In this embodiment, a computer device is provided, such as... Figure 4 As shown, the system includes a memory 1, a processor 2, and a computer program stored in the memory and executable on the processor. The processor executes the computer program to implement the software framework and operating method of the aforementioned coupled simulation method for multiple components and multiple solvers of an aero-engine. Specifically, the computer device can be a computer terminal, a server, or a similar computing device.

[0041] In this embodiment, a computer-readable storage medium is provided, which stores a computer program that executes the above-described multi-component multi-solver coupled simulation method for aero-engines, including a service software framework and a running method.

[0042] Specifically, computer-readable storage media, including both permanent and non-permanent, removable and non-removable media, can store information using any method or technology. Information can be computer-readable instructions, data structures, program modules, or other data. Examples of computer-readable storage media include, but are not limited to, phase-change memory (PRAM), static random access memory (SRAM), dynamic random access memory (DRAM), other types of random access memory (RAM), read-only memory (ROM), electrically erasable programmable read-only memory (EEPROM), flash memory or other memory technologies, CD-ROM, digital versatile optical disc (DVD) or other optical storage, magnetic tape, magnetic disk storage or other magnetic storage devices, or any other non-transferable medium that can be used to store information accessible by a computing device. As defined herein, computer-readable storage media does not include transient computer-readable media, such as modulated data signals and carrier waves.

[0043] Example 2

[0044] See Figure 2 This embodiment uses the initial field setting of a full three-dimensional simulation of a turbofan engine as an example to describe in detail the aero-engine simulation method of the present invention. The engine consists of a three-stage fan, a seven-stage compressor, an annular combustion chamber, a single-stage high-pressure turbine, a single-stage low-pressure turbine, an afterburner, an outer bypass duct, and a nozzle. The specific steps are as follows:

[0045] Step 1: Decomposition and Boundary Condition Setting of Simulation Models for Each Component of the Aero-engine. The engine is decomposed into the fan, compressor, bypass duct, main combustion chamber, high-pressure turbine, low-pressure turbine, afterburner, and nozzle. Simulation models of each component are obtained and meshed. The boundary conditions for the overall aero-engine simulation are set as follows: inlet total temperature of 288K, inlet total pressure of 101325Pa, inlet inflow Mach number of 0.1, and nozzle exit back pressure of 98000Pa.

[0046] Step Two: Based on the flow field characteristics of each component of the aero-engine, match the corresponding turbulence model and solver to the simulation model of each component. Specifically:

[0047] (1) The computational domain of the fan component includes two stages of fans, using a structured grid (approximately 8 million grids), and employing the Spalart-Allmaras (SA) turbulence model in the Reynolds-averaged Navier-Stokes (RANS) equations, with a time step of 1e. -5 s, the solver is a RANS steady-state solver based on the SA turbulence model, and the SA turbulence model parameters and grid point parameters are set;

[0048] (2) The computational domain of the compressor components is the 7-stage compressor blade passage, using a structured grid (structured grid, total number of grids approximately 12 million), and employing the Spalart-Allmaras (SA) turbulence model in the Reynolds-averaged Navier-Stokes (RANS) equations, with a time step set to 1e. -5 s, the solver is a RANS steady-state solver based on the SA turbulence model;

[0049] (3) The computational domain of the main combustion chamber components is the annular combustion chamber (including the flame tube and fuel nozzle), using an unstructured mesh (approximately 50 million meshes). The LES large eddy simulation unsteady solver is used, and a simplified combustion chemical reaction mechanism (containing 12 components and 20 reaction steps) is introduced. The time step is set to 1e. -6 It should be noted that the LES subgrid model of the LES large eddy simulation unsteady solver uses the dynamic Smagorinsky model;

[0050] (4) The computational domain for the high-pressure turbine and the low-pressure turbine is one stage high-pressure turbine + one stage low-pressure turbine. A structured grid (approximately 15 million grids) is used, and a RANS steady-state solver based on the SST turbulence model is employed, with a time step of 1e. -5 s;

[0051] (5) The computational domain of the afterburner is the afterburner (including the igniter and flame stabilizer), using an unstructured mesh (approximately 31 million meshes), and employing the LES large eddy simulation unsteady solver with a time step of 1e. -6 It should be noted that the LES subgrid model of the LES large eddy simulation unsteady solver uses the dynamic Smagorinsky model.

[0052] (6) The computational domain of the nozzle is the convergent-divergent nozzle and exhaust flow field. A structured grid (approximately 6 million grids) is used, and a RANS steady solver based on the SST turbulence model is employed, with a time step of 1e. -5 s.

[0053] This invention employs the Spalart-Allmaras (SA) turbulence model from the Reynolds-averaged Navier-Stokes (RANS) equations for the fan, compressor, and bypass duct components. Leveraging the efficient simulation capabilities of the RANS steady solver for subsonic and transonic turbulence, it reduces computational load while maintaining accuracy. For the turbine component, the SST turbulence model from the RANS steady solver is used. The SST turbulence model's simulation accuracy for high Reynolds number turbulence meets the requirements for turbine flow field calculations.

[0054] Step 3: Based on the boundary conditions of the overall aero-engine simulation, the solvers of each component are driven sequentially according to the airflow sequence of the aero-engine to perform iterative simulation calculations on the flow field parameters between the inlet and outlet of each component. The airflow sequence of the aero-engine is as follows: intake air enters the fan, fan outlet parameters are processed at the interface and then transmitted to the compressor, compressor outlet parameters are transmitted to the main combustion chamber (outer bypass parameters are processed synchronously), combustion chamber outlet parameters are transmitted to the turbine, turbine outlet parameters are transmitted to the afterburner (if afterburner is activated), afterburner outlet parameters are transmitted to the nozzle, and finally exhaust. During the simulation calculation, between two adjacent components, the flow field parameters of the upstream component at the grid points on the interface outlet plane are processed by a radial-circumferential mesh mapping algorithm and then transmitted to the grid points of the downstream component at the interface inlet plane. That is, the flow field parameters output by the solver of the upstream component are processed by the radial-circumferential mesh mapping algorithm at the interface and then output as the inlet boundary conditions of the solver of the downstream component. The solvers for each component perform collaborative calculations at each time step (or iteration step), synchronously iteratively updating the flow field parameters between the inlet and outlet of each component. This continues until the inlet and outlet flow rate deviation of each component is less than a preset flow rate deviation threshold, or the total temperature and total pressure deviation of key inlet and outlet sections is less than a preset deviation threshold. At this point, the overall flow field iteration of the aero-engine is considered to have reached convergence, and the overall simulation results of the aero-engine are output. These simulation results include the thrust value of the aero-engine, the efficiency of each component, the combustion chamber temperature distribution, the flow field distribution between components, and the composition of combustion products. The inlet and outlet flow rate deviation refers to the absolute value of the difference between the inlet and outlet flow rates of a component. The key inlet and outlet sections include the fan inlet and outlet sections, the compressor inlet and outlet sections, and the turbine inlet section. The total temperature and total pressure deviations include total temperature deviation and total pressure deviation, and the corresponding preset deviation thresholds include preset total temperature deviation thresholds and preset total pressure deviation thresholds. For example, the total temperature and total pressure deviations at the fan inlet and outlet sections include the total temperature deviation and total pressure deviation at the fan inlet and outlet sections. The total temperature deviation at the fan inlet and outlet sections refers to the absolute value of the difference between the total temperature at the fan inlet section and the total temperature at the fan outlet section, and the total pressure deviation at the fan inlet and outlet sections refers to the absolute value of the difference between the total pressure at the fan inlet section and the total pressure at the fan outlet section. The total temperature and total pressure deviations at other key inlet and outlet sections are similar. The preset flow rate deviation threshold, preset total temperature deviation threshold, and preset total pressure deviation threshold are all set or adjusted according to simulation needs.

[0055] It should be noted that, between two adjacent components, the flow field parameters of the upstream component at the outlet plane of the interface are transferred to the grid points of the downstream component at the inlet plane of the interface after processing by a radial-circumferential grid mapping algorithm. The methods include:

[0056] Step (1): Determine the interfaces between the various components of the aero-engine, including the interface between the fan and the compressor, the interface between the compressor and the combustion chamber, the interface between the combustion chamber and the turbine, the interface between the turbine and the afterburner, and the interface between the afterburner and the nozzle.

[0057] Step (2): For each interface, the flow field parameters of the upstream component at the grid points on the interface outlet plane are processed using a radial-circumferential grid mapping algorithm and then transferred to the grid points of the downstream component on the interface inlet plane. Specifically, a polar coordinate system is set up on the interface, and the coordinates of the grid points of the upstream component on the interface outlet plane are defined as follows: Define the grid point coordinates of the downstream component on the interface inlet plane as follows: In the grid points of the upstream component on the interface outlet plane, surround the downstream grid points. The coordinates of the four neighboring grid points are respectively , , The following formula is used for bilinear interpolation to map the flow field parameters of the upstream component at the interface outlet plane to the grid points of the downstream component at the interface inlet plane, thereby achieving the transfer of flow field parameters between adjacent components:

[0058] ;

[0059] in: This is an empirical constant, with a value range of 0 to 1. The specific value can be adjusted according to simulation needs or experience. The radial gradient of the flow field parameters at grid points on the inlet plane of the interface of the upstream component; , , , , Grid points , , , , The flow field parameters.

[0060] It should be noted that during the simulation calculation, the flow field parameters of the grid points, including radial velocity, are collected every 10 time steps. Vx Circumferential velocity Vy axial velocity Vz ,temperature T ,pressure P wait.

[0061] Because the radial flow field gradient exists in the region near the endwall of the engine's internal flow channel, the radial-circumferential grid mapping algorithm used in this invention to transfer flow field parameters between adjacent components considers the radial gradient of the flow field parameters at the grid points on the interface outlet plane of the upstream component. This improves the interpolation accuracy in high flow field gradient regions (such as inside the boundary layer) at the engine flow channel end, avoiding the smoothing effect of traditional bilinear interpolation.

[0062] This embodiment employs 24 CPU servers for parallel computation, with a total simulation computation time of 200 hours. When the flow field of the entire system iterates to the 10,000th step, the numerical iteration deviation of the flow field parameters of each component is less than 1e. -5 It reaches a convergent state.

[0063] The above are merely preferred embodiments of the present invention and are not intended to limit the present invention. Any modifications, equivalent substitutions, and improvements made within the spirit and principles of the present invention should be included within the protection scope of the present invention.

Claims

1. A coupled simulation method for multiple components and multiple solvers of an aero-engine, characterized in that, include: Simulation models of each component of the aero-engine are obtained and meshed, and the boundary conditions for the simulation of the entire aero-engine are determined. The components of the aero-engine include the fan, compressor, bypass duct, main combustion chamber, high-pressure turbine, low-pressure turbine, afterburner, and nozzle. The boundary conditions include the total temperature and pressure at the inlet of the aero-engine, the incoming Mach number, and the nozzle exit back pressure. Based on the flow field characteristics of each component of the aero-engine, the corresponding turbulence model and solver are matched for the simulation model of each component; Based on the boundary conditions of the aero-engine whole-machine simulation, a radial-circumferential mesh mapping algorithm is used to transfer the flow field parameters of the grid points on the interface outlet plane of the upstream component to the grid points on the interface inlet plane of the downstream component. The solvers of each component are driven sequentially according to the airflow sequence of the aero-engine to simulate and calculate the flow field within each component of the aero-engine, thereby obtaining the whole-machine simulation results of the aero-engine. The flow field parameters include radial velocity, circumferential velocity, axial velocity, temperature, pressure, and component concentration. The simulation results include the thrust value of the aero-engine, the efficiency of each component, and the combustion chamber temperature distribution. The method for transferring the flow field parameters of the upstream component at the outlet plane of the interface to the grid points of the downstream component at the inlet plane of the interface after processing by a radial-circumferential grid mapping algorithm includes: Step (1): Determine the interfaces between the various components of the aero-engine, including the interface between the fan and the compressor, the interface between the compressor and the combustion chamber, the interface between the combustion chamber and the turbine, the interface between the turbine and the afterburner, and the interface between the afterburner and the nozzle. Step (2): For each interface, the flow field parameters of the upstream component at the grid points on the interface outlet plane are processed by a radial-circumferential grid mapping algorithm and then transferred to the grid points of the downstream component at the interface inlet plane. Specifically, this includes: setting up a polar coordinate system on the interface and defining the coordinates of the grid points of the upstream component at the interface outlet plane as follows: Define the grid point coordinates of the downstream component on the interface inlet plane as follows: In the grid points of the upstream component on the interface outlet plane, surround the downstream grid points. The coordinates of the four neighboring grid points are respectively , , The following formula is used for bilinear interpolation to map the flow field parameters of the upstream component at the interface outlet plane to the grid points of the downstream component at the interface inlet plane, thereby achieving the transfer of flow field parameters between adjacent components: ； in: This is an empirical constant, with a value range of 0 to 1. The specific value can be adjusted according to simulation needs or experience. The radial gradient of the flow field parameters at grid points on the inlet plane of the interface of the upstream component; , , , , Grid points , , , , The flow field parameters.

2. The coupled simulation method for multiple components and multiple solvers of an aero-engine according to claim 1, characterized in that, The solvers of each component are driven sequentially according to the airflow sequence of the aero-engine. When simulating the flow field in each component, the flow field parameters between the inlet and outlet of each component are updated synchronously and iteratively until the inlet and outlet flow deviation of each component is less than the preset flow deviation threshold or the total temperature and total pressure deviation of the key inlet and outlet sections is less than the preset deviation threshold. Then, it is determined that the flow field iteration of the aero-engine has reached the convergence state, and the simulation results of the aero-engine are output.

3. The coupled simulation method for multiple components and multiple solvers of an aero-engine according to claim 1, characterized in that, The fan, compressor, and bypass duct employ the SA turbulence model and a RANS steady-state solver based on the SA turbulence model; the main combustion chamber employs the LES large eddy simulation unsteady solver; the high-pressure turbine, the low-pressure turbine, and the nozzle employ the SST turbulence model and a RANS steady-state solver based on the SST turbulence model; and the afterburner employs the LES large eddy simulation solver.

4. A coupled simulation device for multiple components and multiple solvers of an aero-engine, used to implement the coupled simulation method for multiple components and multiple solvers of an aero-engine as described in any one of claims 1-3, characterized in that, include: The component-level simulation model acquisition module is used to acquire and mesh simulation models of various components of the aero-engine, and simultaneously determine the boundary conditions for the overall simulation of the aero-engine. The components of the aero-engine include the fan, compressor, bypass duct, main combustion chamber, high-pressure turbine, low-pressure turbine, afterburner, and nozzle. The boundary conditions include the total temperature and total pressure at the inlet of the aero-engine, the incoming Mach number, and the nozzle exit back pressure. The turbulence model and solver configuration module is used to match the corresponding turbulence model and solver to the simulation model of each component based on the flow field characteristics of each component of the aero-engine. The simulation execution module, based on the boundary conditions of the overall aero-engine simulation, uses a radial-circumferential mesh mapping algorithm to transfer the flow field parameters of the grid points at the interface outlet plane of the upstream component to the grid points at the interface inlet plane of the downstream component. It sequentially drives the solvers of each component according to the airflow sequence of the aero-engine to perform simulation calculations of the flow field within each component, obtaining the overall simulation results of the aero-engine. The flow field parameters include radial velocity, circumferential velocity, axial velocity, temperature, pressure, and component concentration. The simulation results include the thrust value of the aero-engine, the efficiency of each component, and the combustion chamber temperature distribution. The method for transferring the flow field parameters of the upstream component at the outlet plane of the interface to the grid points of the downstream component at the inlet plane of the interface after processing by a radial-circumferential grid mapping algorithm includes: Step (1): Determine the interfaces between the various components of the aero-engine, including the interface between the fan and the compressor, the interface between the compressor and the combustion chamber, the interface between the combustion chamber and the turbine, the interface between the turbine and the afterburner, and the interface between the afterburner and the nozzle. Step (2): For each interface, the flow field parameters of the upstream component at the grid points on the interface outlet plane are processed by a radial-circumferential grid mapping algorithm and then transferred to the grid points of the downstream component at the interface inlet plane. Specifically, this includes: setting up a polar coordinate system on the interface and defining the coordinates of the grid points of the upstream component at the interface outlet plane as follows: Define the grid point coordinates of the downstream component on the interface inlet plane as follows: In the grid points of the upstream component on the interface outlet plane, surround the downstream grid points. The coordinates of the four neighboring grid points are respectively , , The following formula is used for bilinear interpolation to map the flow field parameters of the upstream component at the interface outlet plane to the grid points of the downstream component at the interface inlet plane, thereby achieving the transfer of flow field parameters between adjacent components: ； in: This is an empirical constant, with a value range of 0 to 1. The specific value can be adjusted according to simulation needs or experience. The radial gradient of the flow field parameters at grid points on the inlet plane of the interface of the upstream component; , , , , Grid points , , , , The flow field parameters.

5. The coupled simulation device for multiple components and multiple solvers of an aero-engine according to claim 4, characterized in that, The solvers of each component are driven sequentially according to the airflow sequence of the aero-engine. When simulating the flow field in each component, the flow field parameters between the inlet and outlet of each component are updated synchronously and iteratively until the inlet and outlet flow deviation of each component is less than the preset flow deviation threshold or the total temperature and total pressure deviation of the key inlet and outlet sections is less than the preset deviation threshold. Then, it is determined that the flow field iteration of the aero-engine has reached the convergence state, and the simulation results of the aero-engine are output.

6. A computer device comprising a memory (1), a processor (2), and a computer program stored in the memory (1) and executable on the processor (2), characterized in that, When the processor (2) executes the computer program, it implements a coupled simulation method for multiple components and multiple solvers of an aero-engine as described in any one of claims 1 to 3.

7. A computer-readable storage medium, characterized in that, The computer-readable storage medium stores a computer program that executes a coupled simulation method for multiple components and multiple solvers of an aero-engine according to any one of claims 1 to 3.

Images