Turbine inlet boundary condition acquisition method, device, equipment and medium

Abstract

The present application relates to the technical field of combustion chamber and turbine, and discloses a turbine inlet boundary condition acquisition method, device, equipment and medium, and provides a turbine inlet boundary condition acquisition method, which comprises the following steps: obtaining input parameters of the combustion chamber, wherein the input parameters at least include geometric parameters, velocity field parameters and temperature field parameters; generating an inlet sector geometry of a current sector based on the geometric parameters; generating an initial velocity field based on the velocity field parameters and the inlet sector geometry, and performing correction processing on the initial velocity field to generate an inlet velocity field; generating an initial temperature field based on the temperature field parameters and the inlet sector geometry, and performing correction processing on the initial temperature field to generate an inlet temperature field; outputting target outlet parameters of the combustion chamber based on the inlet velocity field and the inlet temperature field, and determining the inlet boundary condition of the turbine based on the target outlet parameters. In this way, the defects of large deviation between traditional flow field and temperature field description and difficulty in truly reflecting the non-uniformity of the combustion chamber outlet are effectively overcome.

Description

Technical Field

[0001] This invention relates to the field of combustion chamber and turbine technology, and specifically to a method, apparatus, equipment and medium for obtaining turbine inlet boundary conditions. Background Technology

[0002] The continuous pursuit of high efficiency, low emissions, and high reliability in aero-engines has driven the continuous advancement of combustion technology and turbine aero-thermal design. The application of advanced combustion technology has resulted in a significant non-uniform distribution of aero-thermal parameters at the combustion chamber outlet, mainly manifested as complex flow characteristics such as hot spots (non-uniform temperature field), strong swirling flow, and high turbulence.

[0003] The spatial morphology of hot spots and their circumferential relative position to the turbine inlet guide vanes directly affect the flow field structure and temperature distribution inside the turbine. Unsteady transport and migration of hot spots within the turbine passages and blade tip clearances can lead to deterioration of turbine aerodynamic performance, accelerated blade thermal fatigue, and decreased reliability, posing a severe challenge to turbine aerodynamic design, cooling design, and structural integrity.

[0004] Current methods have significant deviations in describing the characteristics of the internal flow and temperature fields of turbines, making it difficult to accurately reflect the non-uniformity of the combustion chamber outlet and failing to provide accurate and reliable inlet boundary conditions for turbine design. Summary of the Invention

[0005] This invention provides a method, apparatus, equipment, and medium for obtaining turbine inlet boundary conditions, in order to solve the problem of not being able to provide accurate and reliable inlet boundary conditions for turbine design.

[0006] In a first aspect, the present invention provides a method for obtaining turbine inlet boundary conditions, wherein the turbine is disposed downstream of a combustion chamber, the outlet of the combustion chamber is connected to the inlet of the turbine, and the combustion chamber is composed of several sectors; the method includes: Obtain the input parameters of the combustion chamber, which include at least geometric parameters, velocity field parameters, and temperature field parameters; Generate the import sector geometry of the current sector based on the geometric parameters; The initial velocity field is generated based on the velocity field parameters and the geometry of the inlet sector, and the initial velocity field is then corrected to generate the inlet velocity field. The initial temperature field is generated based on the temperature field parameters and the geometry of the inlet sector, and the initial temperature field is then corrected to generate the inlet temperature field. Based on the inlet velocity field and inlet temperature field, the target outlet parameters of the combustion chamber are output, and the inlet boundary conditions of the turbine are determined based on the target outlet parameters.

[0007] In this embodiment, by acquiring the geometric parameters, velocity field parameters, and temperature field parameters of the combustion chamber, the inlet sector geometry is generated. Based on this, the initial velocity field and initial temperature field are progressively corrected to obtain a high-precision inlet velocity field and inlet temperature field. This allows for the determination of the turbine inlet boundary conditions, which can accurately characterize the non-uniform gas-heat distribution characteristics such as the hot spot morphology, strong swirling flow, and high turbulence at the combustion chamber outlet. This effectively overcomes the shortcomings of traditional flow field and temperature field descriptions, which have large deviations and fail to accurately reflect the non-uniformity of the combustion chamber outlet. Therefore, accurate and reliable inlet boundary conditions are provided for turbine aerodynamic design, cooling design, and structural integrity assessment, helping to reduce the risk of blade thermal fatigue and improve turbine reliability and overall engine performance.

[0008] In one optional implementation, the geometric parameters include at least the axial coordinates of the current sector, the initial values ​​of the circumferential coordinates, the number of turbine inlet guide vane blades, and the flow channel curve; the flow channel curve includes the upper endwall flow channel curve and the lower endwall flow channel curve. Based on geometric parameters, generate the inbound sector geometry for the current sector, including: Based on the axial coordinate, interpolation calculations are performed on the upper and lower endwall flow channel curves to obtain the radial coordinate range. Calculate the sector angle based on the number of turbine inlet guide vane blades; The range of circumferential coordinates is determined based on the sector angle and the initial value of the circumferential coordinates. The boundary region of the current sector is determined based on the radial coordinate range and the circumferential coordinate range. The number of radial and circumferential grid points is determined based on the radial coordinate range and sector angle. The geometry of the inlet sector is determined based on the boundary region, the number of radial grid points, and the number of circumferential grid points.

[0009] In this embodiment, on the one hand, by performing interpolation calculations on the flow channel curves of the upper and lower end walls and combining them with axial coordinates, the radial coordinate range can be accurately obtained, achieving high-precision geometric reconstruction of the current sector boundary region and ensuring the accuracy of the physical model. On the other hand, by determining the number of grid points based on the radial coordinate range and sector angle, the generated grid can adapt to the geometric characteristics of the sector, avoiding grid distortion and providing a high-quality computational grid foundation for subsequent flow field and temperature field calculations.

[0010] In one alternative implementation, an initial velocity field is generated based on velocity field parameters and inlet sector geometry, including: Within the geometrically defined boundary region of the inlet sector, the initial velocity field induced by a single vortex core is calculated based on a preset turbine model.

[0011] In this embodiment, by introducing a preset turbine model and calculating the initial velocity field induced by a single vortex core, the strong swirling flow and complex vortex structure at the combustion chamber outlet can be accurately simulated, solving the problem that traditional methods are difficult to capture high turbulence and swirling characteristics, and significantly improving the realism of the velocity field simulation.

[0012] In one alternative implementation, the initial velocity field is corrected, including: The initial velocity field is sequentially modified by periodic correction, endwall constraint correction, and near-wall swirl rebalancing correction.

[0013] In one optional implementation, the velocity field parameters include at least the swirling direction, swirling intensity, and vortex core diameter; applying periodic corrections, endwall constraint corrections, and near-wall swirling rebalancing corrections sequentially to the initial velocity field includes: False vortex cores are arranged inside the turbine inlet guide vane passages of each adjacent sector of the current sector. The velocity fields induced by each false vortex core are superimposed on the velocity field of the current sector so that the velocity distribution on the boundary of the adjacent sector satisfies the periodic boundary condition. The false vortex cores have the same swirling direction, swirling intensity and vortex core diameter as the vortex core of the current sector. The initial velocity field after periodic correction is sequentially corrected by mirror spurious vortex correction, endwall tilt angle correction, and mirror spurious vortex circumferential velocity correction. The circumferential velocity correction is constructed according to the preset exponential decay function; Based on the circumferential velocity correction, the circumferential swirling velocity in the near-lower endwall and near-upper endwall regions of the current sector is corrected respectively.

[0014] In this embodiment, on the one hand, by arranging spurious vortex cores with consistent parameters in adjacent sectors and superimposing their induced velocity fields, periodic matching of the velocity distribution at the boundary of adjacent sectors is achieved, eliminating abrupt changes in the flow field at the periodic boundary. On the other hand, by sequentially performing mirror spurious vortex correction, endwall inclination angle correction, and circumferential velocity correction, and combining this with an exponential decay function to construct a near-wall correction for targeted adjustment, the swirling attenuation characteristics near the endwall can be accurately captured. This effectively restores the true strong swirling and high turbulence characteristics at the combustion chamber outlet, solving the problem of large flow field description deviations in the prior art. This provides high-fidelity inlet boundary conditions for turbine design, helps to accurately predict hot spot migration trajectories, and improves turbine aerodynamic performance and reliability.

[0015] In one alternative implementation, the initial temperature field is corrected, including: The temperature correction field of the cold air film and the strip-shaped temperature correction field are sequentially superimposed onto the initial temperature field.

[0016] In this embodiment, by sequentially superimposing the cold gas film temperature correction field and the strip-shaped temperature correction field, the complex non-uniform temperature field formed at the outlet of the advanced combustion chamber due to the mixing of cooling airflow and non-uniform combustion can be accurately simulated. This truly reflects the spatial morphology and local temperature gradient of the hot spot, effectively improving the accuracy of the description of the hot spot characteristics at the combustion chamber outlet and providing more reliable temperature boundary conditions for turbine cooling design and thermal analysis.

[0017] In an alternative implementation, before outputting the target outlet parameters of the combustion chamber based on the inlet velocity field and the inlet temperature field, the method further includes: Visual verification was performed on the inlet velocity field and the inlet temperature field, respectively. If the inlet velocity field or inlet temperature field does not meet the preset requirements, adjust the input parameters and regenerate the inlet velocity field and inlet temperature field. If both the inlet velocity field and the inlet temperature field meet the preset requirements, output the target outlet parameters.

[0018] In this embodiment, by introducing a visualization verification step, it is possible to intuitively and effectively verify whether the generated inlet velocity field and temperature field truly reflect the non-uniform characteristics such as hot spots and strong swirl at the combustion chamber outlet, thus ensuring the physical reliability of the target outlet parameters.

[0019] Secondly, the present invention provides a turbine inlet boundary condition acquisition device, wherein the turbine is disposed downstream of the combustion chamber, the outlet of the combustion chamber is connected to the inlet of the turbine, and the combustion chamber is composed of several sectors. The device includes: The acquisition module is used to acquire the input parameters of the combustion chamber, which include at least geometric parameters, velocity field parameters, and temperature field parameters. The generation module is used to generate the inlet sector geometry of the current sector based on geometric parameters; The first correction module is used to generate an initial velocity field based on velocity field parameters and inlet sector geometry, and to correct the initial velocity field based on the velocity field parameters to generate the inlet velocity field. The second correction module is used to generate an initial temperature field based on temperature field parameters and inlet sector geometry, and to correct the initial temperature field based on the temperature field parameters to generate the inlet temperature field. The output module is used to output the target outlet parameters of the combustion chamber based on the inlet velocity field and the inlet temperature field, and to determine the inlet boundary conditions of the turbine based on the target outlet parameters.

[0020] Thirdly, the present invention provides an electronic device, comprising: a memory and a processor, wherein the memory and the processor are communicatively connected to each other, the memory stores computer instructions, and the processor executes the computer instructions to perform the turbine inlet boundary condition acquisition method of the first aspect or any corresponding embodiment described above.

[0021] Fourthly, the present invention provides a computer-readable storage medium storing computer instructions for causing a computer to execute the turbine inlet boundary condition acquisition method of the first aspect or any corresponding embodiment thereof. Attached Figure Description

[0022] To more clearly illustrate the specific embodiments of the present invention or the technical solutions in the prior art, the drawings used in the description of the specific embodiments or the prior art will be briefly introduced below. Obviously, the drawings described below are some embodiments of the present invention. For those skilled in the art, other drawings can be obtained from these drawings without creative effort.

[0023] Figure 1 This is a schematic flowchart of the first method for obtaining turbine inlet boundary conditions according to an embodiment of the present invention; Figure 2 This is a schematic diagram of the second process for obtaining turbine inlet boundary conditions according to an embodiment of the present invention; Figure 3 This is a schematic diagram of the velocity component induced by a single Lamb-Oseen vortex core; Figure 4 This is a schematic diagram of the periodic correction of sectors; Figure 5 This is a schematic diagram of the vortex velocity field in a mirror. Figure 6 This is a schematic diagram of the velocity field component caused by the endwall inclination angle; Figure 7 This is a schematic diagram of the circumferential velocity correction of the mirror vortex; Figure 8 This is a schematic diagram of near-wall swirl correction; Figure 9 This is a schematic diagram of the temperature distribution of the initial temperature field; Figure 10 This is a schematic diagram of the cooling film temperature correction. Figure 11 This is a schematic diagram of strip-type temperature correction; Figure 12 It is the distribution of temperature and velocity in the current sector generated by the mathematical model of the temperature field; Figure 13 This is a structural block diagram of a turbine inlet boundary condition acquisition device according to an embodiment of the present invention; Figure 14 This is a schematic diagram of the hardware structure of an electronic device according to an embodiment of the present invention. Detailed Implementation

[0024] To make the objectives, technical solutions, and advantages of the embodiments of the present invention clearer, the technical solutions of the embodiments of the present invention will be clearly and completely described below with reference to the accompanying drawings. Obviously, the described embodiments are only some embodiments of the present invention, not all embodiments. Based on the embodiments of the present invention, all other embodiments obtained by those skilled in the art without creative effort are within the scope of protection of the present invention.

[0025] It is understood that before using the technical solutions disclosed in the various embodiments of the present invention, users should be informed of the types, scope of use, and usage scenarios of the personal information involved in the present invention and their authorization should be obtained in accordance with relevant laws and regulations through appropriate means.

[0026] The terms "first" and "second" are used for descriptive purposes only and should not be construed as indicating or implying relative importance or implicitly specifying the number of technical features indicated. Thus, a feature defined as "first" or "second" may explicitly or implicitly include one or more of that feature. In the description of this invention, "a plurality of" means two or more, unless otherwise explicitly specified.

[0027] The combustion chamber outlet flow field represents the incoming flow boundary conditions at the turbine inlet, where velocity and temperature distribution characteristics directly influence the flow, heat transfer, and cooling design of turbine components. However, due to the complex flow, mixing, and combustion processes within the combustion chamber, the temperature and velocity fields at the outlet exhibit significant circumferential and radial non-uniform distributions. The outlet temperature field contains "hot spots" with high-temperature cores, which, combined with residual swirling flow, create strong non-uniform intake distortion. Studies show that this non-uniform inflow significantly affects the flow structure and heat transfer characteristics within the turbine blade cascade. Ignoring the actual distortion characteristics at the combustion chamber outlet during turbine design will lead to inaccuracies in turbine thermal load prediction.

[0028] Currently, the methods for obtaining inlet boundary conditions in turbine aerodynamic design and thermal analysis mainly suffer from the following problems: First, traditional design methods oversimplify turbine inlet boundary conditions. In conventional independent engine component design processes, turbine inlet boundary conditions are typically based on one-dimensional lumped parameters or two-dimensional circumferential average distributions, failing to reflect the three-dimensional non-uniform characteristics of the hot spot and swirling flow at the combustion chamber exit. This simplification of boundary conditions, deviating from the actual combustion chamber-turbine coupled flow field in a real engine, directly affects the consistency between predicted turbine aero-thermal performance and overall engine testing.

[0029] Second, current parametric design methods for combustion chambers extract geometric parameters and mesh parameters as variables to achieve automated simulation. However, while aiming to optimize the combustion chamber geometry to improve the outlet temperature distribution coefficient, they do not establish a parametric mathematical model for the non-uniform distribution of the combustion chamber outlet velocity and temperature fields. For turbine design, there is still a lack of a parametric tool for inlet boundary parameters that can independently generate and flexibly control hot spots and swirl characteristics.

[0030] Third, in recent years, large eddy simulation and adaptive turbulence simulation methods have been used for combustion chamber-turbine coupled simulation, which can capture the hot spot and residual swirling structure at the combustion chamber exit and their impact on the turbine. However, these high-precision simulation methods require huge computational resources and have long cycles, making them difficult to widely apply in the initial iterative stage of turbine design. In engineering practice, it is necessary to have an efficient and controllable parametric means in the component design stage to generate turbine inlet boundary conditions that reflect the real physical characteristics.

[0031] According to an embodiment of the present invention, a method for obtaining turbine inlet boundary conditions is provided. It should be noted that the steps shown in the flowchart in the accompanying drawings can be executed in a computer system such as a set of computer-executable instructions. Furthermore, although a logical order is shown in the flowchart, in some cases, the steps shown or described may be executed in a different order than that shown here.

[0032] This embodiment provides a method for obtaining turbine inlet boundary conditions, which can be used in the aforementioned electronic device. The turbine is located downstream of the combustion chamber, and the outlet of the combustion chamber is connected to the inlet of the turbine. The combustion chamber consists of several sectors. Figure 1 This is a schematic flowchart of the first method for obtaining turbine inlet boundary conditions according to an embodiment of the present invention, as shown below. Figure 1 As shown, the process includes the following steps: Step S101: Obtain the input parameters of the combustion chamber. The input parameters include at least geometric parameters, velocity field parameters, and temperature field parameters.

[0033] Here, the combustion chamber refers to the component in an aircraft engine that converts the chemical energy of fuel into thermal energy, and its outlet airflow directly enters the turbine inlet.

[0034] Input parameters refer to various control variables preset by the user, including geometric parameters used to determine the shape of the sector, velocity field parameters used to control the generation of swirl, and temperature field parameters used to control the generation of hot spots.

[0035] The geometric parameters include at least the axial position of the current sector, the circumferential starting angle, the number of turbine inlet guide vanes, and the flow channel curves of the upper and lower endwalls. They are mainly used to determine the spatial range and grid distribution of the computational domain.

[0036] The velocity field parameters include at least the axial intake velocity, the relative position of the vortex core center, the vortex direction, the vortex intensity and the sampling range of the vortex core diameter, as well as the attenuation factor range of each correction element. They are mainly used to control the intensity, size and position of the residual vortex at the combustion chamber outlet.

[0037] The temperature field parameters include at least the average total temperature of the inlet section, the total temperature non-uniformity factor, the total temperature and thickness of the end-wall film cooling, the circumferential non-uniformity factor of the hot spot, and the range of attenuation factors for each temperature correction element. These parameters are mainly used to control the shape and intensity of the hot spot at the combustion chamber outlet.

[0038] It should be noted that the entire annular combustion chamber outlet section is divided into several circumferentially divided sector regions according to the number of turbine inlet guide vane blades. Each sector corresponds to a turbine inlet guide vane pitch (i.e., a guide vane channel), which serves as the basic calculation unit for subsequent parametric modeling.

[0039] Step S102: Generate the import sector geometry of the current sector based on the geometric parameters.

[0040] Here, the import sector geometry refers to the annular sector region and its discrete grid within the current sector area, bounded by the radial boundary (upper and lower end walls) and the circumferential boundary (center lines of adjacent guide vanes on the left and right sides). It is the spatial region that needs to be satisfied for the calculation of the velocity field and temperature field.

[0041] Step S103: Generate an initial velocity field based on the velocity field parameters and the geometry of the inlet sector, and then correct the initial velocity field to generate the inlet velocity field.

[0042] Here, the initial temperature field refers to the initial swirling velocity distribution induced by a single vortex core generated on the sector geometry based on the Lamb-Oseen vortex model, used to simulate the characteristics of the annular high-speed region and the central low-speed region of the residual swirling flow at the outlet of a real combustion chamber.

[0043] It should be noted that, since the initial velocity field does not satisfy the physical constraints of the turbine flow field, periodic correction, endwall constraint correction and near-wall swirl rebalancing correction need to be applied sequentially to ensure that the final velocity field satisfies the periodic conditions on the boundaries of adjacent sectors, the non-penetration condition at the endwall, and the characteristics of the annular high-speed region near the swirl center.

[0044] Step S104: Generate an initial temperature field based on temperature field parameters and inlet sector geometry, and then correct the initial temperature field to generate the inlet temperature field.

[0045] Here, the initial temperature field refers to the temperature field of the heat dissipation spot generated by superimposing circumferential sinusoidal disturbances on the radial parabolic distribution of the total inlet temperature, reflecting the high-temperature core and circumferential non-uniformity of the hot spot at the combustion chamber outlet.

[0046] It should be noted that since the initial temperature field does not take into account the end-wall film cooling effect and the average temperature deviation, end-wall film cooling correction and strip-shaped average temperature compensation correction need to be applied sequentially to ensure that the final temperature field has a near-wall low temperature layer and maintains a constant global average total temperature.

[0047] Step S105: Based on the inlet velocity field and inlet temperature field, output the target outlet parameters of the combustion chamber, and determine the inlet boundary conditions of the turbine based on the target outlet parameters.

[0048] Here, the target outlet parameters include at least the coordinate point, total temperature, and airflow angles in three directions.

[0049] In this step, the corrected velocity field (converted to airflow angles in three directions) and temperature field data are written into a standard format file (such as a user configuration file in CSV (Comma-Separated Values) format). This file can be directly imported into CFD (Computational Fluid Dynamics) software as turbine inlet boundary conditions.

[0050] In this embodiment, by acquiring the geometric parameters, velocity field parameters, and temperature field parameters of the combustion chamber, the inlet sector geometry is generated. Based on this, the initial velocity field and initial temperature field are progressively corrected to obtain a high-precision inlet velocity field and inlet temperature field. This allows for the determination of the turbine inlet boundary conditions, which can accurately characterize the non-uniform gas-heat distribution features such as the hot spot morphology, strong swirling flow, and high turbulence at the combustion chamber outlet. This effectively overcomes the shortcomings of existing technologies, such as large deviations in flow field and temperature field descriptions and the inability to accurately reflect the non-uniformity of the combustion chamber outlet. Therefore, accurate and reliable inlet boundary conditions are provided for turbine aerodynamic design, cooling design, and structural integrity assessment, helping to reduce the risk of blade thermal fatigue and improve turbine reliability and overall engine performance.

[0051] This embodiment provides a method for obtaining turbine inlet boundary conditions, which can be used in the aforementioned electronic equipment. Figure 2 This is a schematic diagram of the second process for obtaining turbine inlet boundary conditions according to an embodiment of the present invention, as shown below. Figure 2 As shown, the process includes the following steps: Step S201: Obtain the input parameters of the combustion chamber. The input parameters include at least geometric parameters, velocity field parameters, and temperature field parameters. For details, please refer to [link to relevant documentation]. Figure 1 Step S101 of the illustrated embodiment will not be described again here.

[0052] Step S202: Generate the import sector geometry of the current sector based on the geometric parameters.

[0053] In some optional implementations, the geometric parameters include at least the axial coordinates of the current sector, the initial values ​​of the circumferential coordinates, the number of turbine inlet guide vane blades, and the flow channel curve; the flow channel curve includes the upper endwall flow channel curve and the lower endwall flow channel curve, and the above step S202 includes: Step a1: Based on the axial coordinates, interpolation calculations are performed on the upper and lower endwall flow channel curves to obtain the radial coordinate range. Step a2: The sector angle is calculated based on the number of turbine inlet guide vane blades. Step a3: The circumferential coordinate range is determined based on the sector angle and the initial circumferential coordinate value. Step a4: The boundary region of the current sector is determined based on the radial and circumferential coordinate ranges. Step a5: The number of radial and circumferential grid points is determined based on the radial coordinate range and the sector angle. Step a6: The inlet sector geometry is determined based on the boundary region, the number of radial grid points, and the number of circumferential grid points.

[0054] Here, the upper endwall flow channel curve and the lower endwall flow channel curve refer to the axially varying geometric shapes of the outer wall (casing side) and inner wall (hub side) of the turbine inlet annular flow channel, respectively, and are usually input in the form of discrete points or spline curves.

[0055] The radial coordinate range refers to the radial inner and outer boundaries obtained by interpolating the flow channel curves of the upper and lower end walls at a given axial position.

[0056] The number of turbine inlet guide vane blades refers to the number of blades in the first-stage turbine inlet guide vane, which is used to determine the total number of sectors into which the annular cross section is divided.

[0057] The sector angle refers to the circumferential angle occupied by each turbine inlet guide vane passage.

[0058] The starting value of the circumferential coordinates refers to the starting angle of the circumferential direction of the current sector, which is used to determine the specific position of the sector in the entire annular cross section.

[0059] The circumferential coordinate range refers to the circumferential boundary of the current sector, which is determined by the starting angle and the sector angle. That is, the interval from the starting angle to the starting angle plus the sector angle.

[0060] The radial grid point count refers to the number of computational nodes arranged along the radial direction, which is preset by the user according to the accuracy requirements and determines the radial resolution.

[0061] The number of circumferential grid points refers to the calculated number of nodes arranged along the circumferential direction. It is determined by the ratio of arc length to development height and is used to ensure that the physical dimensions of the grid cells are relatively balanced in the circumferential and radial directions.

[0062] Import sector geometry refers to the complete computational domain containing boundary regions and grid discretization information, which is the spatial basis for the generation of the initial velocity field and the initial temperature field.

[0063] It should be noted that the specific steps for obtaining the geometry of the inlet sector are as follows: 1. Based on the input axial coordinate value, interpolation (usually linear or spline interpolation) is performed on the stored upper and lower endwall flow channel curve data to extract the hub radius and casing radius at that axial position, thus obtaining the radial coordinate range. The interpolation accuracy depends on the data density of the flow channel curve and the choice of interpolation method. 2. By dividing the whole circumference by the number of turbine inlet guide vane blades, the sector angle corresponding to each guide vane pitch is obtained. 3. Based on the input circumferential coordinate starting value, the calculated sector angle is added to obtain the circumferential ending value, thus determining the complete circumferential range of the current sector from start to end. 4. The obtained radial coordinate range and circumferential coordinate range are combined to form an annular sector enclosed by the radial inner and outer boundaries and the circumferential left and right boundaries, which is the physical spatial boundary of the current sector. 5. By reading the user-preset radial grid point count, and then automatically calculating the circumferential grid point count according to the sector angle, average radius, and spanwise height, based on the arc length to spanwise height ratio, the physical dimensions of the grid cells in the circumferential and radial directions are matched. 6. Within the defined boundary region, computational nodes are evenly distributed according to the defined radial and circumferential grid point numbers to form a complete structured computational grid, which serves as the spatial discretization basis for the subsequent generation of velocity and temperature fields.

[0064] In some implementations, the axial coordinate values ​​of the inlet sector are known. The maximum radial coordinate of the MP40 section can be obtained by interpolation in the upper and lower endwall flow channel curves. and minimum value .

[0065] Given the number of turbine inlet guide vanes Circumferential coordinate starting value Find the sector angle of the imported sector. for Then the circumferential coordinate termination value of the imported sector for .

[0066] Therefore, the radial coordinate range of the sector is known. and circumferential coordinate range This allows us to determine the geometric extent of the current sector. Given the radial grid point count of the inlet sector, the circumferential grid point count is approximately determined by the ratio of the radial grid point count to the circumferential grid point count, which is roughly equal to the ratio between the arc length and the elevation (the circumferential and radial coordinates of a point in the sector are represented by...). and (represented by the following formula):

[0067] in, The number of radial grid points. The number of points in the circumferential grid. For sector angle, The maximum value of the radial coordinate. This is the minimum value of the radial coordinate.

[0068] In this embodiment, on the one hand, by performing interpolation calculations on the flow channel curves of the upper and lower end walls and combining them with axial coordinates, the radial coordinate range can be accurately obtained, achieving high-precision geometric reconstruction of the current sector boundary region and ensuring the accuracy of the physical model. On the other hand, by determining the number of grid points based on the radial coordinate range and sector angle, the generated grid can adapt to the geometric characteristics of the sector, avoiding grid distortion and providing a high-quality computational grid foundation for subsequent flow field and temperature field calculations.

[0069] Step S203: Generate an initial velocity field based on the velocity field parameters and the geometry of the inlet sector, and then correct the initial velocity field to generate the inlet velocity field.

[0070] Specifically, step S203 includes: Step S2031: Within the boundary region defined by the geometry of the inlet sector, the initial velocity field induced by a single vortex core is calculated based on a preset turbine model.

[0071] Here, the preset turbine model refers to the Lamb-Oseen vortex model, which is used to simulate the velocity distribution of the residual swirling flow at the combustion chamber exit. This model can generate a ring-shaped velocity field rotating around the center of the vortex core, with the velocity at the center being zero, gradually increasing outward to a peak value and then slowly decaying.

[0072] It should be noted that the specific steps for obtaining the initial velocity field are as follows: Calculate the distance from the node to the vortex core center: Based on the actual physical coordinates of the node and the vortex core center, calculate the Euclidean distance L to calculate the induced velocity magnitude. Substitute the distance L into the Lamb-Oseen vortex velocity formula to obtain the induced velocity magnitude at that node. This velocity magnitude exhibits a distribution characteristic of first increasing and then decreasing from the vortex core center outwards. Based on the input swirl direction parameters (clockwise or counterclockwise), determine the direction of the induced velocity to be perpendicular to the line connecting the node and the vortex core center. Repeat the above calculation for all nodes within the mesh to generate a complete initial velocity field, which is then saved as a vector map or contour map.

[0073] In some alternative implementations, in a Cartesian coordinate system, the coordinates of the vortex core center of the Lamb-Oseen vortex are set as follows: The coordinates of a certain point are The distance between this point and the center of the vortex core is defined by the following formula:

[0074] Then the Lamb-Oseen vortex is at point Induction velocity value at the location for:

[0075] The induced velocity direction is perpendicular to the line connecting this point and the center of the vortex core, expressed as the vortex direction. Defined, this value can take ±1, representing the clockwise or counterclockwise vortex velocity respectively. The swirling intensity represents the vortex strength; the stronger the swirling intensity, the greater the velocity at a point in the induced velocity field. This indicates the diameter of the vortex core.

[0076] The relative circumferential coordinates are defined by the following formula. and relative radial coordinates :

[0077] Given the relative circumferential coordinates of the vortex core center and relative radial coordinates This allows us to determine the location of the vortex core center within the current sector.

[0078] The relative vortex core diameter is defined by the following formula. :

[0079] This characterizes the proportion of the turbine inlet vortex core diameter to the current sector spanwise height.

[0080] To further illustrate, the component velocity field induced by a single Lamb-Oseen vortex core is as follows: Figure 3 As shown, Figure 3 This is a schematic diagram of the velocity component induced by a single Lamb-Oseen vortex core.

[0081] In the diagram, the cloud map represents the magnitude of the velocity value, the black arrows indicate the direction and relative magnitude of the velocity, and the solid black dots represent the location of the vortex core center. Observing the flow within the sector in the above figure, we can see that: the flow velocity near the vortex core center is very low, and the flow velocity increases with distance from the vortex core center; a ring-shaped high-speed zone surrounds the vortex core center, where the flow velocity reaches its peak; outside the ring-shaped high-speed zone, the flow velocity decreases with distance from the vortex core center. Looking at the flow field near the sector boundary, we can see that the velocity component induced by a single vortex core does not satisfy the periodicity condition of the turbine flow field at the boundary of adjacent sectors, nor does it satisfy the physical constraint of "no airflow in or out" at the endwall. Therefore, it is necessary to correct the velocity component field induced by a single vortex core. The correction method is to superimpose the periodically corrected velocity component field of the sector and the mirror-image illusory vortex velocity component field.

[0082] In this embodiment, by introducing a preset turbine model and calculating the initial velocity field induced by a single vortex core, the strong swirling flow and complex vortex structure at the combustion chamber outlet can be accurately simulated, solving the problem that traditional methods are difficult to capture high turbulence and swirling characteristics, and significantly improving the realism of the velocity field simulation.

[0083] In some optional implementations, step S203 above further includes: Step S2032: Apply periodic correction, endwall constraint correction and near-wall swirl rebalancing correction to the initial velocity field in sequence.

[0084] In some optional implementations, the velocity field parameters include at least the swirling direction, swirling intensity, and vortex core diameter; step S2032 further includes: Step b1: Arrange false vortex cores inside the turbine inlet guide vane passages of each adjacent sector of the current sector, and superimpose the velocity fields induced by each false vortex core into the velocity field of the current sector so that the velocity distribution on the boundary of the adjacent sector satisfies the periodic boundary condition. The false vortex cores and the vortex cores of the current sector have the same swirling direction, swirling intensity and vortex core diameter.

[0085] It should be noted that the velocity field induced by the Lamb-Oseen vortex is corrected based on the turbine flow field characteristics to obtain the turbine inlet velocity field distribution, thereby calculating the turbine inlet airflow angle distribution.

[0086] Turbine inlet velocity field It is composed of the superposition of various component velocity fields, including the component velocity field induced by a single Lamb-Oseen vortex core. Periodically corrected velocity field of sector Mirror vortex velocity field The velocity field component caused by the end wall inclination angle Circumferential velocity correction component of mirror vortex and near-wall swirling correction of the velocity field component .Right now:

[0087] Specifically, since the velocity field induced by a single Lamb-Oseen vortex core does not satisfy the periodic boundary conditions at the interface between adjacent sectors, periodic sector correction is required. This means that, except for the current sector, a "false vortex" is placed in each turbine inlet guide vane passage. The vortex core's swirling direction, swirling intensity, vortex core diameter, and relative position within its respective passage are exactly the same as the vortex core described earlier.

[0088] in, Let be the circumferential coordinates of the i-th illusory vortex core. These are the circumferential coordinates of the actual vortex core within the current sector. For the i-th sector angle, Let be the radial coordinate of the i-th illusory vortex core. The radial coordinates of the actual vortex core within the current sector. Let i be the swirling direction of the i-th spurious vortex core. This represents the swirling direction of the actual vortex core within the current sector. Let be the swirl intensity of the i-th spurious vortex core. The swirling intensity of the actual vortex core, Let be the relative vortex diameter of the i-th illusory vortex core. This represents the relative vortex core diameter of the current sector.

[0089] It is possible Figure 4 As shown, Figure 4 This is a schematic diagram of the periodic correction of the sector. Figure 4 Figure (a) shows the sector-periodic correction velocity field, and Figure (b) shows the initial velocity field of the superimposed sector-periodic correction velocity field.

[0090] As shown in Figure (a), the periodically corrected velocity field mainly operates at the sector boundaries, making the airflow magnitude and angle approximately the same on the left and right boundaries of the sector. In other words, the airflow magnitude and angle between adjacent channels are approximately the same. As shown in Figure (b), within the sector, the velocity component used for periodic correction is approximately the circumferential velocity, which weakens the circumferential velocity of the swirling flow below the mid-width and strengthens the circumferential velocity of the swirling flow above the mid-width.

[0091] It should be noted that Figure (a) contains the formula for the sector periodically corrected velocity field, and Figure (b) contains the formula for the sector periodically corrected velocity field superimposed with the vortex core-induced velocity field.

[0092] Step b2 involves sequentially performing mirror spurious vortex correction, endwall tilt angle correction, and mirror spurious vortex circumferential velocity correction on the initial velocity field after periodic correction.

[0093] To further explain, for each vortex core within the channel, "mirror-like false vortices" are superimposed on the lower and upper end walls, respectively, using the lower and upper end walls as symmetry planes.

[0094] in, Let H be the circumferential position of the illusory vortex core reflected by the i-th endwall mirror, with subscripts H and S indicating the lower endwall (Hub) and upper endwall (Shroud), respectively. Let represent the swirling direction of the illusory vortex core reflected by the i-th endwall mirror. Let be the swirling intensity of the illusory vortex core reflected by the i-th endwall mirror surface. Let be the relative vortex diameter of the illusory vortex cores reflected by the i end-wall mirrors.

[0095] The radial coordinates of the mirror vortex core with the endwall as the plane of symmetry below are different from those of the mirror vortex core with the endwall as the plane of symmetry above, as follows:

[0096] in, Let be the radial coordinates of the illusory vortex core reflected by the mirror surface of the i-th endwall, which is symmetrical about the following endwalls. Let be the radial coordinates of the illusory vortex core reflected by the mirror surface of the i-th endwall with the upper endwall as the plane of symmetry.

[0097] Specifically, such as Figure 5 As shown, Figure 5 This is a schematic diagram of the velocity field of the mirror vortex component. Figure 5 In this context, (a) represents the velocity component induced by the mirror vortex core with the endwall as the plane of symmetry, and includes its formula explanation; Figure 5 In the equation (b), the velocity component induced by the mirror vortex core with the upper endwall as the plane of symmetry is given, and its formula is explained. Figure 5 In the equation (c), the initial velocity field after superimposing the mirror vortex velocity field is included, along with its formula explanation.

[0098] That is, the velocity field induced by the symmetrical mirror vortex core of the lower end wall acts throughout the entire fan-shaped area, mainly by modifying the airflow angle near the lower end wall and the circumferential velocity throughout the entire fan-shaped area; the velocity field induced by the symmetrical mirror vortex core of the upper end wall acts near the upper end wall, mainly by modifying the airflow angle near the upper end wall, but has little effect on the circumferential velocity throughout the entire fan-shaped area. (Observation) Figure 5 As shown in (a), the circumferential velocity component induced by the mirror vortex core is very large throughout the entire fan, which has a significant impact on the distribution of the airflow angle and needs to be corrected again. The method of correction is to correct the circumferential velocity component field of the mirror vortex.

[0099] Ideally, the airflow near the endwall should adhere to the endwall, and the airflow angle should equal the endwall tilt angle; the farther away from the endwall, the less the airflow is affected by the endwall tilt angle. A normalized attenuation function is introduced.

[0100]

[0101] In the above formula, K is the attenuation factor. The larger the value of the attenuation factor, the greater the function f. The faster it decays; t For normalized coordinates, the coordinates and normalization method need to be selected according to the direction of decay.

[0102] Let the inclination angles of the upper and lower end walls at section MP40 be respectively and The endwall inclination angle will result in a radial component on this cross-section. Let the estimated axial intake velocity be... Then, by the normalized decay function f The defined second-order flow velocity field is:

[0103] in, For the secondary flow component velocity field, The endwall inclination angle at section MP40. This is the attenuation factor for the effect of the endwall tilt angle. These are normalized distance coordinates.

[0104] The normalized coordinates selected in the above formula are as follows:

[0105] in, Normalized distance coordinates representing the influence of the lower end wall inclination angle. is the normalized distance coordinate influenced by the inclination angle of the upper wall, and r is the radial coordinate of the current sector.

[0106] Specifically, such as Figure 6 As shown, Figure 6 A schematic diagram of the velocity field component resulting from the endwall inclination angle. Figure 6 In this context, (a) represents the velocity field component caused by the inclination angle of the lower end wall, and includes its formula explanation; Figure 6 In the diagram, (b) represents the velocity field component caused by the inclination angle of the upper end wall, and includes its formula explanation; Figure 6 (c) in the figure represents the initial velocity field after the velocity field component is generated by the superposition of the endwall inclination angle, and includes its formula explanation.

[0107] observe Figure 6 As shown in (a) and (b), the lower end wall has a larger inclination angle, so the radial velocity component caused by the inclination angle of the lower end wall is also larger, reaching a maximum of 14.63 m / s. The farther away from the lower end wall, the smaller the radial velocity component caused by the inclination angle of the lower end wall. The upper end wall has a smaller inclination angle, so the radial velocity component caused by the inclination angle of the upper end wall is very small, with a maximum radial velocity of only 2.59 m / s.

[0108] Step b3: Construct the circumferential velocity correction amount according to the preset exponential decay function.

[0109] Furthermore, when a mirror vortex is superimposed on the endwall as a symmetry plane, a circumferential velocity component is superimposed on most locations within the current sector. This results in a "full-circumference" swirling flow on the MP40 section, which is detrimental to the study of the swirling flow within the sector. To counteract the negative impact of the mirror vortex, a circumferential velocity correction is needed for the current velocity field. The circumferential velocity correction component is defined as:

[0110] in, The velocity field is corrected for the circumferential velocity. The circumferential velocity is induced by the mirror vortex core with the lower endwall as the plane of symmetry.

[0111] The circumferential velocity (a full-circumference vortex) induced by the mirror vortex core with the endwalls as symmetry planes is calculated using the above formula. Subtracting this circumferential velocity from the current velocity field is the process of circumferential velocity correction. The velocity field after circumferential velocity correction is as follows: Figure 7 As shown.

[0112] in Figure 7 In the figure, (a) represents the velocity component field corrected for the circumferential velocity of the mirror vortex, and includes its formula explanation; Figure 7 (b) in the figure represents the velocity field after superimposing the velocity component corrected for the circumferential velocity of the mirror vortex, and includes its formula explanation.

[0113] It should be noted that in Figure (a), the minimum superimposed circumferential velocity correction value is approximately 35.92 m / s, the maximum superimposed circumferential velocity correction value is approximately 44.91 m / s, and the average value is approximately 39.98 m / s. In Figure (b), a high-speed flow region exists above the swirl center, but the flow velocity is relatively low below the swirl center. Based on the velocity vector distribution characteristics at the combustion chamber outlet section, the swirl velocity is approximately the same at positions symmetrical about the swirl center point in the spanwise direction. Therefore, the flow field characteristics in Figure (b) still deviate significantly from the velocity vector distribution characteristics at the combustion chamber outlet measurement section and require correction.

[0114] Step b4: Based on the circumferential velocity correction, the circumferential swirling velocity in the near-lower endwall and near-upper endwall regions of the current sector is corrected respectively.

[0115] It should be noted that the velocity field induced by a single Lamb-Oseen vortex core does not satisfy many constraints of the turbine flow field. Although the corrected velocity field can satisfy the constraints of the turbine flow field, it loses important "vortex flow characteristics". For example, the swirling velocity near the lower endwall almost disappears, the swirling velocity near the upper endwall is too large, and the annular high-velocity region around the center of the vortex core disappears.

[0116] Therefore, it is necessary to correct the initial velocity field through near-wall swirl correction, that is, to amplify the excessively small circumferential swirl velocity and reduce the excessively large circumferential swirl velocity, so that the corrected velocity field is as close as possible to the velocity field induced by the single Lamb-Oseen vortex core while satisfying the constraints of the turbine flow field.

[0117] The near-wall swirling correction velocity component is defined by the attenuation function. Ideally, the circumferential swirling velocity in the near-wall region should be close to the circumferential swirling velocity induced by a single Lamb-Oseen vortex nucleus in the near-wall region. Therefore, the swirling correction velocity components near the lower and upper endwalls are defined as follows:

[0118] in, The swirling correction velocity components are for the near-lower and near-upper endwalls. The circumferential velocity induced by a single Lamb-Oseen vortex core in the near-wall region. This represents the average circumferential velocity in the near-wall region of the velocity field. The attenuation factor is a correction for near-wall swirl flow. Normalized distance coordinates corrected for near-wall swirl.

[0119] like Figure 8 As shown, Figure 8 This is a schematic diagram of the near-wall swirling flow correction. Figure 8 In this context, (a) represents the swirling corrected velocity field near the lower endwall, and includes its formula explanation; Figure 8 (b) in the figure represents the swirling corrected velocity field near the upper endwall, and includes its formula explanation; Figure 8 (c) in the figure represents the initial velocity field after superimposing the near-wall swirling correction velocity field, and includes its formula explanation.

[0120] Depend on Figure 8 It can be seen that after near-wall swirl correction, the annular high-velocity region around the vortex core reappears, and the excessively small or large circumferential swirl velocity is corrected, and the characteristics of the velocity field are more consistent with the velocity field distribution characteristics on the measurement section at the combustion chamber outlet.

[0121] In this embodiment, on the one hand, by arranging spurious vortex cores with consistent parameters in adjacent sectors and superimposing their induced velocity fields, periodic matching of the velocity distribution at the boundary of adjacent sectors is achieved, eliminating abrupt changes in the flow field at the periodic boundary. On the other hand, by sequentially performing mirror spurious vortex correction, endwall inclination angle correction, and circumferential velocity correction, and combining this with an exponential decay function to construct a near-wall correction for targeted adjustment, the swirling attenuation characteristics near the endwall can be accurately captured. This effectively restores the true strong swirling and high turbulence characteristics at the combustion chamber outlet, solving the problem of large flow field description deviations in the prior art. This provides high-fidelity inlet boundary conditions for turbine design, helps to accurately predict hot spot migration trajectories, and improves turbine aerodynamic performance and reliability.

[0122] Step S204: Generate an initial temperature field based on temperature field parameters and inlet sector geometry, and then correct the initial temperature field to generate the inlet temperature field.

[0123] In some alternative implementations, step S204 further includes step c1.

[0124] Step c1: The temperature correction field of the cold air film and the strip-shaped temperature correction field are sequentially superimposed onto the initial temperature field.

[0125] In this step, the temperature field of a single heat dissipation spot (i.e., the initial temperature field) is corrected based on the total temperature distribution characteristics of the turbine flow field to obtain the turbine inlet temperature field distribution.

[0126] The turbine inlet temperature field T is composed of the temperature field of a single heat dissipation spot. It is formed by superimposing the temperature correction field, which includes the temperature correction field of the cold air film. and strip-shaped temperature correction field ,Right now:

[0127] It should be noted that the temperature field away from the heat dissipation spot is obtained by correcting the radial distribution of the inlet total temperature. The radial distribution of the inlet total temperature is a parabolic function. Function expression:

[0128] The coefficients can be obtained by fitting the experimental data. , and The value of r is the radial coordinate of the current sector, and the radial coordinates corresponding to the vertex of the parabola are also calculated. As the distance from the center of the heat dissipation spot, the temperature distribution at that distance is:

[0129] in, This is the circumferential non-uniformity coefficient; the larger the value, the more drastic the circumferential temperature change. The average total temperature at the turbine inlet section is determined by the design parameters provided by the client. The initial phase determines the circumferential position of the hot spot center in the current sector; This refers to the temperature field away from the heat dissipation spot; The total inlet temperature is linearly distributed radially. The temperature value at the peak radius of the radial parabolic distribution; The total temperature is linearly distributed at the vertex radius; This represents the number of blades in the high-pressure guide vane.

[0130] Theoretically, the temperature distribution of the heat sink is adjusted relative to the radial distribution of the total inlet temperature by a correction factor, which is related to the radial coordinate. and circumferential coordinates The trigonometric functions, the magnitude of which depends only on the radial coordinate. This is relevant. If we integrate this correction over the entire circumference of the turbine inlet section, we will find that the integral result of the correction is zero. That is, correcting the radial distribution of the inlet total temperature using the above method will not change the average total temperature of the turbine inlet section.

[0131] like Figure 9 As shown, Figure 9 This is a schematic diagram of the initial temperature field. The contour plot represents the magnitude of the temperature values, and the solid magenta dots indicate the locations of the hotspot centers. In the temperature distribution field away from the heat sink, the total temperature near the upper endwall is approximately 1742 K, and the total temperature near the lower endwall is approximately 1452 K. Observing the above figure, it can be seen that the temperature distribution field away from the heat sink cannot simulate the influence of the film cooling on the endwall surface on the total temperature distribution across the MP40 cross-section. The temperature distribution field away from the heat sink needs to be corrected by superimposing a temperature correction field from the cold film. The center point of the high-temperature region in the temperature distribution field away from the heat sink coincides with the parabola. radial coordinates of the vertex There are also some discrepancies, which need to be corrected to make the two overlap as much as possible.

[0132] Furthermore, by considering the effects of the cold air film on the lower end wall surface and the upper end wall respectively, the air film temperature of the upper end wall is... The air film temperature at the lower end wall is approximately 1000 K. It is approximately 1200 K. Theoretically, the cooling film is a thin film attached to the end wall surface and does not affect the mainstream temperature field. At the same time, it can be assumed that within the thickness of the film, due to the mixing with the mainstream combustion gas, the further away from the end wall surface, the worse the cooling effect and the higher the temperature.

[0133] The temperature correction field of the cooling film is defined by the decay function:

[0134] in, For the temperature correction field of the cold air film, The total temperature of the end-wall air film. The attenuation factor is used to correct for the temperature of the cooling film. Normalized distance coordinates corrected for the cooling film.

[0135] The method for selecting the normalized coordinates in the above formula is as follows:

[0136] in, and These represent the relative thicknesses of the cooling air film on the lower and upper end walls, respectively. Relative thickness refers to the proportion of the air film thickness to the sector's expansion height. Normalized distance coordinates corrected for the lower end wall cooling film. Normalized distance coordinates corrected for the cooling film on the upper wall.

[0137] Specifically, such as Figure 10 As shown, Figure 10 (a) in the figure represents the temperature correction field of the cold air film on the lower end wall, and its temperature is 1200K. Figure 10 (b) in the figure represents the temperature correction field of the upper end wall cold air film, and its temperature is 1000 K. Figure 10 (c) in the figure represents the initial temperature field after the cold air film temperature correction field is superimposed, and includes its formula explanation.

[0138] Among them, through Figure 10 It is known that the temperature correction range of the cold air film is limited to the vicinity of the end wall surface. Slightly further away from the end wall surface, the corrected temperature of the cold air film quickly decays to zero, and the temperature correction of the cold air film does not change the temperature distribution characteristics within the sector. However, due to the presence of the temperature correction field of the cold air film, the average temperature across the entire sector is reduced, and its value no longer matches the provided design parameters. Therefore, it is necessary to correct the temperature field after superimposing the temperature correction field of the cold air film. The correction method is to superimpose a strip-shaped temperature correction field.

[0139] Furthermore, since this temperature correction field only changes in the radial direction, the high-temperature region shown on the contour map resembles a strip, hence it is called a strip-shaped temperature correction field.

[0140] The purpose of superimposing the strip-shaped temperature correction field is twofold: firstly, to correct the average total temperature in the figure to be consistent with the provided design parameters; and secondly, to correct the center point of the hot spot in the figure to be near the center of the high-temperature region.

[0141] The strip-shaped temperature correction field is defined by the decay function:

[0142] in, It is a strip-type temperature correction field. The peak temperature of the strip-shaped temperature correction field. The attenuation factor is a temperature-corrected banding factor. Normalized distance coordinates for strip-shaped temperature correction.

[0143] The method for selecting the normalized coordinates in the above formula is as follows:

[0144] in, Normalized distance coordinates for strip-shaped temperature correction.

[0145] Specifically, such as Figure 11 As shown, Figure 11 This is a schematic diagram of strip-type temperature correction. Figure 11 (a) in the figure represents the strip-type temperature correction field. Figure 11 (b) in the figure represents the temperature field after superimposing the strip-type temperature correction field, and includes its formula explanation.

[0146] As shown in the figure, the highest temperature of the strip-shaped temperature correction field is 35.36 K, and the corrected average total temperature of the fan-shaped area is 1699.23 K, which differs from the average total temperature of the turbine inlet under design conditions by 25 K, with a relative error of only about -1.5%. Furthermore, the radial coordinates of the high-temperature zone in Figure (a) are related to the parabola... radial coordinates of the vertex By overlapping, the high-temperature area in Figure (b) is pulled to a higher position, so that the center point of the high-temperature area coincides with the parabola. radial coordinates of the vertex The gap between them has narrowed.

[0147] In this embodiment, by sequentially superimposing the cold gas film temperature correction field and the strip-shaped temperature correction field, the complex non-uniform temperature field formed at the outlet of the advanced combustion chamber due to the mixing of cooling airflow and non-uniform combustion can be accurately simulated. This truly reflects the spatial morphology and local temperature gradient of the hot spot, effectively improving the accuracy of the description of the hot spot characteristics at the combustion chamber outlet and providing more reliable temperature boundary conditions for turbine cooling design and thermal analysis.

[0148] Step S205: Based on the inlet velocity field and inlet temperature field, output the target outlet parameters of the combustion chamber, and determine the inlet boundary conditions of the turbine based on the target outlet parameters.

[0149] In some implementations, specifically as follows Figure 12 As shown, Figure 12The temperature and velocity distributions generated for the mathematical model of the temperature field in the current sector, where, Figure 12 (a) in the figure is a temperature distribution map, including the average temperature. The value is 1699.2329K, and the relative error δ is -1.4937%. Figure 12 (b) in the diagram shows the temperature and velocity distribution, including the highest temperature value. The K value is 1967.775 K, and the relative error δ is 3.5435%. As shown in the figure, the hot spot center is located on the sector boundary, and the vortex center is located in the center of the sector. The circumferential position difference between the hot spot center and the vortex center is significant. This aligns with the characteristic that "the center point of the highest temperature zone and the swirling center point will differ by half a fan angle in the circumferential direction." Under design conditions, the hot spot temperature on the MP40 section is approximately 1900.43 K, and the highest temperature value generated by the model within the current fan is 1967.775 K, resulting in a temperature difference of 67.34 K and a relative error of only 3.5435%.

[0150] In summary, in the temperature distribution field generated by the mathematical model of the temperature field, the relative error between the average temperature and the design operating condition value is only -1.5%, and the relative error between the hot spot temperature and the design operating condition value is only 3.5%. The temperature distribution characteristics are very similar to the total temperature distribution characteristics on the measurement section, and it can ensure that the relationship between the flow field and temperature field generated by the parameterized model conforms to the relationship between the flow field and temperature field on the measurement section.

[0151] In some alternative implementations, step S205 may be preceded by: Step d1: Visually verify the inlet velocity field and inlet temperature field respectively. Step d2: If the inlet velocity field or inlet temperature field does not meet the preset requirements, adjust the input parameters and regenerate the inlet velocity field and inlet temperature field. Step d3: If both the inlet velocity field and inlet temperature field meet the preset requirements, output the target outlet parameters.

[0152] Here, the preset requirements refer to the pre-set physical rationality judgment criteria, including but not limited to: whether the velocity field meets the periodic boundary conditions (the velocity distribution on the left and right boundaries is consistent), whether the velocity near the end wall meets the non-penetration condition, whether the velocity field has the characteristics of a ring-shaped high-speed region around the vortex core center, whether the average total temperature of the temperature field is consistent with the design value, and whether the relative position of the hot spot center and the vortex center conforms to physical laws, etc.

[0153] In some implementations, after generating the inlet velocity field and the inlet temperature field, a velocity vector distribution map (displaying the direction and magnitude of the velocity at each node), a velocity magnitude cloud map (using color depth to represent velocity amplitude), an isotherm distribution map (displaying temperature contour lines), and a temperature magnitude cloud map (using color to represent temperature level) can be generated by calling a drawing function.

[0154] By observing the above images, it can be determined whether the velocity field has characteristics such as a ring-shaped high-speed region, a low-speed region at the center of the vortex, and whether the velocities at the left and right boundaries are symmetrical and continuous. At the same time, it can be determined whether the temperature field has characteristics such as a hot spot high-temperature core, a low-temperature layer at the end wall, and a circumferential temperature gradient.

[0155] In some implementations, if visualization verification reveals that the physical field does not meet the preset requirements (such as significant inconsistency in the velocities at the left and right boundaries of the velocity field, or excessive deviation of the average temperature of the temperature field from the design value), the corresponding input parameters are adjusted after analyzing the cause of the deviation (such as modifying the position of the vortex core center, adjusting the attenuation factor value, or changing the air film temperature), and the program is rerun to generate a new inlet velocity field and inlet temperature field. Visual verification is then performed again until the requirements are met.

[0156] In some implementations, after the inlet velocity field and inlet temperature field have been verified by visualization, all grid nodes are traversed, and the coordinates, total temperature and airflow angles in three directions of each node are extracted. An output file is written in standard CSV format, and this output file can be used as the turbine inlet boundary condition.

[0157] In this embodiment, by introducing a visualization verification step, it is possible to intuitively and effectively verify whether the generated inlet velocity field and temperature field truly reflect the non-uniform characteristics such as hot spots and strong swirl at the combustion chamber outlet, thus ensuring the physical reliability of the target outlet parameters.

[0158] This embodiment also provides a turbine inlet boundary condition acquisition device, which is used to implement the above embodiments and preferred embodiments; details already described will not be repeated. As used below, the term "module" can refer to a combination of software and / or hardware that performs a predetermined function. Although the device described in the following embodiments is preferably implemented in software, hardware implementation, or a combination of software and hardware, is also possible and contemplated.

[0159] This embodiment provides a device for obtaining turbine inlet boundary conditions, such as... Figure 13 As shown, the turbine is located downstream of the combustion chamber, and the outlet of the combustion chamber is connected to the inlet of the turbine. The combustion chamber consists of several sectors, and the device includes: The acquisition module 1301 is used to acquire the input parameters of the combustion chamber, which include at least geometric parameters, velocity field parameters, and temperature field parameters.

[0160] The generation module 1302 is used to generate the import sector geometry of the current sector based on geometric parameters.

[0161] The first correction module 1303 is used to generate an initial velocity field based on velocity field parameters and inlet sector geometry, and to correct the initial velocity field based on the velocity field parameters to generate an inlet velocity field.

[0162] The second correction module 1304 is used to generate an initial temperature field based on temperature field parameters and inlet sector geometry, and to correct the initial temperature field based on the temperature field parameters to generate an inlet temperature field.

[0163] The output module 1305 is used to output the target outlet parameters of the combustion chamber based on the inlet velocity field and the inlet temperature field, and to determine the inlet boundary conditions of the turbine based on the target outlet parameters.

[0164] In some optional implementations, the geometric parameters include at least the axial coordinates of the current sector, the initial values ​​of the circumferential coordinates, the number of turbine inlet guide vane blades, and the flow channel curve; the flow channel curve includes the upper endwall flow channel curve and the lower endwall flow channel curve. Generation module 1302 includes: The first calculation submodule is used to perform interpolation calculations on the upper and lower endwall flow channel curves based on the axial coordinates to obtain the radial coordinate range.

[0165] The second calculation submodule is used to calculate the sector angle based on the number of turbine inlet guide vane blades.

[0166] The first determination submodule is used to determine the range of circumferential coordinates based on the sector angle and the initial value of the circumferential coordinates.

[0167] The second determination submodule is used to determine the boundary region of the current sector based on the radial coordinate range and the circumferential coordinate range.

[0168] The third determination submodule is used to determine the number of radial grid points and the number of circumferential grid points based on the radial coordinate range and sector angle.

[0169] The fourth determination submodule is used to determine the geometry of the inlet sector based on the boundary region, the number of radial grid points, and the number of circumferential grid points.

[0170] In some alternative implementations, the first correction module 1303 includes: The third calculation submodule is used to calculate the initial velocity field induced by a single vortex core within the boundary region defined by the geometry of the inlet sector, based on a preset turbine model.

[0171] In some alternative implementations, the first correction module 1303 includes: The first correction submodule is used to sequentially apply periodic correction, endwall constraint correction, and near-wall swirl rebalancing correction to the initial velocity field.

[0172] In some optional implementations, the velocity field parameters include at least the swirling direction, swirling intensity, and vortex core diameter; the first correction submodule includes: The superposition unit is used to arrange false vortex cores inside the turbine inlet guide vane passages of each adjacent sector of the current sector, and superimpose the velocity fields induced by each false vortex core onto the velocity field of the current sector so that the velocity distribution on the boundary of the adjacent sector satisfies the periodic boundary conditions. The false vortex cores have the same swirling direction, swirling intensity and vortex core diameter as the vortex core of the current sector. The first correction unit is used to sequentially perform mirror false vortex correction, endwall tilt angle correction, and mirror false vortex circumferential velocity correction on the initial velocity field.

[0173] The construction unit is used to construct the circumferential velocity correction according to a preset exponential decay function.

[0174] The second correction unit is used to correct the circumferential swirling velocity of the near-lower endwall and near-upper endwall regions of the current sector based on the circumferential velocity correction amount.

[0175] In some alternative implementations, the second correction module 1304 includes: The superposition submodule is used to sequentially superimpose the temperature correction field of the cold air film and the strip-shaped temperature correction field onto the initial temperature field.

[0176] In some alternative implementations, the method further includes, before outputting the target outlet parameters of the combustion chamber based on the inlet velocity field and the inlet temperature field: The verification module is used to perform visual verification of the inlet velocity field and the inlet temperature field, respectively.

[0177] The adjustment module is used to adjust the input parameters and regenerate the inlet velocity field and inlet temperature field when the inlet velocity field or inlet temperature field does not meet the preset requirements.

[0178] The module is used to output target outlet parameters when both the inlet velocity field and the inlet temperature field meet preset requirements.

[0179] The turbine inlet boundary condition acquisition device provided in this embodiment of the invention can execute the turbine inlet boundary condition acquisition method provided in any embodiment of the invention, and has the corresponding functional modules and beneficial effects of the method. Further functional descriptions of the above modules and units are the same as in the corresponding embodiments described above, and will not be repeated here.

[0180] Figure 14 This is a schematic diagram of the structure of an electronic device provided in an embodiment of the present invention.

[0181] The following is a detailed reference. Figure 14 The diagram illustrates a structural schematic suitable for implementing an electronic device according to embodiments of the present invention. The electronic device may include a processor (e.g., a central processing unit, graphics processor, etc.) 1401, which can perform various appropriate actions and processes according to a program stored in read-only memory (ROM) 1402 or a program loaded from memory 1408 into random access memory (RAM) 1403. The RAM 1403 also stores various programs and data required for the operation of the electronic device. The processor 1401, ROM 1402, and RAM 1403 are interconnected via a bus 1404. An input / output (I / O) interface 1405 is also connected to the bus 1404.

[0182] Typically, the following devices can be connected to I / O interface 1405: input devices 1406 including, for example, touchscreens, touchpads, keyboards, mice, cameras, microphones, accelerometers, gyroscopes, etc.; output devices 1407 including, for example, liquid crystal displays (LCDs), speakers, vibrators, etc.; memory devices 1408 including, for example, magnetic tapes, hard disks, etc.; and communication devices 1409. Communication device 1409 allows electronic devices to communicate wirelessly or wiredly with other devices to exchange data. Although Figure 14 Electronic devices with various devices are shown, but it should be understood that it is not required to implement or have all of the devices shown, and more or fewer devices may be implemented or have instead.

[0183] In particular, according to embodiments of the present invention, the processes described above with reference to the flowcharts can be implemented as computer software programs. For example, embodiments of the present invention include a computer program product comprising a computer program carried on a non-transitory computer-readable medium, the computer program containing program code for performing the methods shown in the flowcharts. In such embodiments, the computer program can be downloaded and installed from a network via a communication device 1409, or installed from a memory 1408, or installed from a ROM 1402. When the computer program is executed by the processor 1401, it performs the functions defined in the turbine inlet boundary condition acquisition method of the embodiments of the present invention.

[0184] Figure 14 The electronic device shown is merely an example and should not be construed as limiting the functionality and scope of use of the embodiments of the present invention.

[0185] This invention also provides a computer-readable storage medium. The methods described above according to embodiments of the invention can be implemented in hardware or firmware, or implemented as computer code that can be recorded on a storage medium, or implemented as computer code downloaded via a network and originally stored on a remote storage medium or a non-transitory machine-readable storage medium and then stored on a local storage medium. Thus, the methods described herein can be processed by software stored on a storage medium using a general-purpose computer, a dedicated processor, or programmable or dedicated hardware. The storage medium can be a magnetic disk, optical disk, read-only memory, random access memory, flash memory, hard disk, or solid-state drive, etc.; further, the storage medium can also include combinations of the above types of memory. It is understood that computers, processors, microprocessor controllers, or programmable hardware include storage components capable of storing or receiving software or computer code. When the software or computer code is accessed and executed by the computer, processor, or hardware, the turbine inlet boundary condition acquisition method shown in the above embodiments is implemented.

[0186] A portion of this invention can be applied as a computer program product, such as computer program instructions, which, when executed by a computer, can invoke or provide the methods and / or technical solutions according to the invention through the operation of the computer. Those skilled in the art will understand that the forms in which computer program instructions exist in a computer-readable medium include, but are not limited to, source files, executable files, installation package files, etc. Correspondingly, the ways in which computer program instructions are executed by a computer include, but are not limited to: the computer directly executing the instructions, or the computer compiling the instructions and then executing the corresponding compiled program, or the computer reading and executing the instructions, or the computer reading and installing the instructions and then executing the corresponding installed program. Here, the computer-readable medium can be any available computer-readable storage medium or communication medium accessible to a computer.

[0187] Although embodiments of the invention have been described in conjunction with the accompanying drawings, those skilled in the art can make various modifications and variations without departing from the spirit and scope of the invention, and such modifications and variations all fall within the scope defined by the appended claims.

Claims

1. A method for obtaining turbine inlet boundary conditions, characterized in that, The turbine is located downstream of the combustion chamber, the outlet of the combustion chamber is connected to the inlet of the turbine, and the combustion chamber consists of several sectors. The method includes: The input parameters of the combustion chamber are obtained, and the input parameters include at least geometric parameters, velocity field parameters, and temperature field parameters; Based on the geometric parameters, generate the inlet sector geometry for the current sector; An initial velocity field is generated based on the velocity field parameters and the geometry of the inlet sector, and the initial velocity field is then corrected to generate the inlet velocity field. An initial temperature field is generated based on the temperature field parameters and the geometry of the inlet sector, and the initial temperature field is then corrected to generate the inlet temperature field. Based on the inlet velocity field and the inlet temperature field, the target outlet parameters of the combustion chamber are output, and the inlet boundary conditions of the turbine are determined based on the target outlet parameters.

2. The method according to claim 1, characterized in that, The geometric parameters include at least the axial coordinates, initial circumferential coordinates, number of turbine inlet guide vane blades, and flow channel curve of the current sector; the flow channel curve includes the upper endwall flow channel curve and the lower endwall flow channel curve. The step of generating the inlet sector geometry of the current sector based on the geometric parameters includes: Based on the axial coordinates, interpolation calculations are performed on the upper end wall flow channel curve and the lower end wall flow channel curve respectively to obtain the radial coordinate range; Calculate the sector angle based on the number of turbine inlet guide vane blades; The circumferential coordinate range is determined based on the sector angle and the initial value of the circumferential coordinates. The boundary region of the current sector is determined based on the radial coordinate range and the circumferential coordinate range; The number of radial grid points and the number of circumferential grid points are determined based on the radial coordinate range and the sector angle; The geometry of the inlet sector is determined based on the boundary region, the number of radial grid points, and the number of circumferential grid points.

3. The method according to claim 1, characterized in that, The generation of the initial velocity field based on the velocity field parameters and the inlet sector geometry includes: Within the boundary region defined by the geometry of the inlet sector, the initial velocity field induced by a single vortex core is calculated based on a preset turbine model.

4. The method according to claim 3, characterized in that, The correction process for the initial velocity field includes: The initial velocity field is sequentially subjected to periodic correction, endwall constraint correction, and near-wall swirl rebalancing correction.

5. The method according to claim 4, characterized in that, The velocity field parameters include at least the swirling direction, swirling intensity, and vortex core diameter; the sequential application of periodic correction, endwall constraint correction, and near-wall swirling rebalancing correction to the initial velocity field includes: False vortex cores are arranged inside the turbine inlet guide vane passages of each adjacent sector of the current sector. The velocity fields induced by each false vortex core are superimposed on the velocity field of the current sector so that the velocity distribution on the boundary of the adjacent sector satisfies the periodic boundary condition. The false vortex cores have the same swirling direction, swirling intensity and vortex core diameter as the vortex core of the current sector. The initial velocity field after periodic correction is sequentially corrected by mirror spurious vortex correction, endwall tilt angle correction, and mirror spurious vortex circumferential velocity correction. The circumferential velocity correction is constructed according to the preset exponential decay function; Based on the circumferential velocity correction, the circumferential swirling velocities of the near-lower endwall and near-upper endwall regions of the current sector are corrected respectively.

6. The method according to claim 1, characterized in that, The correction process for the initial temperature field includes: The temperature correction field of the cold air film and the strip-shaped temperature correction field are sequentially superimposed onto the initial temperature field.

7. The method according to claim 1, characterized in that, Before outputting the target outlet parameters of the combustion chamber based on the inlet velocity field and the inlet temperature field, the method further includes: The inlet velocity field and the inlet temperature field were visually verified respectively. If the inlet velocity field or the inlet temperature field does not meet the preset requirements, adjust the input parameters and regenerate the inlet velocity field and the inlet temperature field. If both the inlet velocity field and the inlet temperature field meet the preset requirements, the target outlet parameters are output.

8. A device for obtaining turbine inlet boundary conditions, characterized in that, The turbine is located downstream of the combustion chamber, and the outlet of the combustion chamber is connected to the inlet of the turbine. The combustion chamber consists of several sectors. The device includes: The acquisition module is used to acquire the input parameters of the combustion chamber, which include at least geometric parameters, velocity field parameters, and temperature field parameters; A generation module is used to generate the inlet sector geometry of the current sector based on the geometric parameters; The first correction module is used to generate an initial velocity field based on the velocity field parameters and the geometry of the inlet sector, and to correct the initial velocity field based on the velocity field parameters to generate the inlet velocity field. The second correction module is used to generate an initial temperature field based on the temperature field parameters and the geometry of the inlet sector, and to correct the initial temperature field based on the temperature field parameters to generate the inlet temperature field. The output module is used to output the target outlet parameters of the combustion chamber based on the inlet velocity field and the inlet temperature field, and to determine the inlet boundary conditions of the turbine based on the target outlet parameters.

9. An electronic device, characterized in that, include: A memory and a processor are communicatively connected, the memory stores computer instructions, and the processor executes the turbine inlet boundary condition acquisition method according to any one of claims 1 to 7 by executing the computer instructions.

10. A computer-readable storage medium, characterized in that, The computer-readable storage medium stores computer instructions for causing a computer to perform the turbine inlet boundary condition acquisition method according to any one of claims 1 to 7.

Images