Heavy-duty gas turbine aero-configuration design method and system, electronic device and storage medium

Abstract

The application provides a heavy gas turbine air-cooled turbine aerodynamic layout design method and system, electronic equipment and a storage medium. The method comprises the following steps: step 1: obtaining the design requirements of the heavy gas turbine air-cooled turbine, and determining the turbine stage number; step 2: determining the enthalpy drop and load distribution scheme of each turbine stage; step 3: performing preliminary evaluation on the stress of the blade and the disc, if the requirement is not met, adjusting the load coefficient and the flow coefficient, if the requirement is met, entering step 4; step 4: initial meridian flow passage design, evaluating the strength of the blade and the disc after the meridian flow passage design is completed, if the requirement is not met, repeating steps 2 to 4 to perform parameter optimization, if the requirement is met, entering step 5; step 5: determining the radial distribution law of the turbine stage reaction degree; step 6: determining the radial distribution law of the aerodynamic parameter according to the one-dimensional average parameter design result; step 7: S2 inverse problem calculation; and step 8: aerodynamic performance analysis and optimization. The scheme provided by the application shortens the turbine engineering design cycle.

Description

TECHNICAL FIELD

[0001] The present application relates to the technical field of heavy-duty gas turbine design, and particularly relates to a heavy-duty gas turbine gas-cooled turbine aerodynamic layout design method and system, an electronic device and a storage medium. BACKGROUND

[0002] With the gradual increase of the power level of the gas turbine and the inlet gas temperature of the turbine, the turbine aerodynamic load and thermal load are continuously improved. The increase of the power level of the gas turbine leads to the increase of the turbine flow and the blade load, the size of the turbine last-stage blade is larger, and the blade and the disc are more likely to encounter problems in strength and vibration. Therefore, at the aerodynamic design stage, the constraints of the blade and the disc in strength and vibration need to be considered, and corresponding optimization adjustment needs to be performed.

[0003] As the upstream of the turbine component design work, the selection of the configuration, the flow passage, the parameter and the layout in the turbine aerodynamic low-dimensional design plays a great role in the difficulty and rationality of the subsequent cooling, structure, strength and the like. However, the design parameters mainly focused on by the turbine at the low-dimensional aerodynamic design stage at present mainly include the power and distribution of the turbine stage, the reaction degree, the flow coefficient, the load coefficient, the flow angle and the Mach number, and the influence of the aerodynamic layout on the cooling, the strength and vibration, the structural integrity and the like is less considered.

[0004] The existing turbine design system mainly compensates for the lag of the aerodynamic design result relative to the professional needs of cooling, structure, strength and the like through repeated iteration between different professionals inside the turbine and between the turbine and the overall structure. That is, after completing a complete design iteration, the adjustment of the aerodynamic low-dimensional design scheme is considered according to the structural strength evaluation result, so that the aerodynamic efficiency is ensured, and the professional needs of cooling, structure, strength and the like are met. This inevitably leads to the problems of many iteration times, long cycle, large workload and repetition of the turbine aerodynamic design. Therefore, it is urgent to propose a scheme for quickly designing the aerodynamic layout of the heavy-duty gas turbine gas-cooled turbine. SUMMARY

[0005] The present application provides a heavy-duty gas turbine gas-cooled turbine aerodynamic layout design method, system, electronic device and storage medium to at least solve the technical problems of many iteration times, long cycle, large workload and repetition of the turbine aerodynamic design.

[0006] The first aspect embodiment of the present application provides a heavy-duty gas turbine gas-cooled turbine aerodynamic layout design method, which comprises:

[0007] Step 1: obtaining the design requirements of the heavy-duty gas turbine gas-cooled turbine, determining the design input parameters of the turbine based on the design requirements, and determining the turbine stage number according to the design criteria, wherein the design input parameters comprise: inlet gas flow, total pressure, total temperature, power, efficiency, cold gas flow and flow passage basic size;

[0008] Step 2: Load distribution of turbine stage, based on cooling requirement of downstream blade and strength constraint of blade, determine enthalpy drop and load distribution scheme of each turbine stage;

[0009] Step 3: One-dimensional design parameter selection, based on the load distribution result of stage, select load coefficient, flow coefficient and reaction degree of each stage of turbine, and calculate the product of blade exit annular area and square of gas turbine speed and tip and rim speed, preliminarily evaluate the stress of blade and disc through simplified model, if the strength requirement is not met, adjust the load coefficient and flow coefficient, if met, enter step 4;

[0010] Step 4: Primary meridian flow passage design, based on aerodynamic parameters and overall structure size constraint, determine flow path layout, turbine diameter and blade height, blade axial width and axial spacing, meridian flow passage expansion angle, complete meridian flow passage design, and evaluate blade and disc strength, if not meet the requirement, repeat step 2 to step 4 for parameter optimization, if meet, enter step 5;

[0011] Step 5: Flow pattern design, based on controllable vortex design method of controlling blade row exit circulation, determine the radial distribution law of turbine stage reaction degree;

[0012] Step 6: Radial distribution law design of aerodynamic parameters, according to the design result of one-dimensional average parameter, determine the radial distribution law of aerodynamic parameters;

[0013] Step 7: S2 inverse problem calculation, through multi-iteration calculation of S2 flow surface, obtain velocity triangle, aerodynamic thermodynamic parameters and detailed performance parameters of each radial section of turbine stage;

[0014] Step 8: Aerodynamic performance analysis and optimization, optimize average aerodynamic parameters and radial distribution law, repeat step 3 to step 8 until the aerodynamic layout design result of heavy-duty gas turbine air-cooled turbine meeting the requirements of aerodynamic performance and strength is obtained.

[0015] Preferably, the determination of enthalpy drop and load distribution scheme of each turbine stage based on cooling requirement of downstream blade and strength constraint of blade comprises:

[0016] Based on the cooling requirement of downstream blade and the strength constraint of blade, and using step-by-step descending load distribution form, the enthalpy drop and load distribution scheme of each stage is determined.

[0017] Further, the selection of load coefficient, flow coefficient and reaction degree of each stage of turbine based on the load distribution result of stage comprises:

[0018] The load coefficient of each stage of turbine is selected based on the size of turbine diameter and turbine stage efficiency;

[0019] The flow coefficient is selected based on the turbine blade height and the exit Mach number limit of the cascade.

[0020] Selecting reaction degree based on outlet Mach number of moving blade, axial force and cooling design boundary condition.

[0021] Further, the preliminary evaluation of the stress of the blade and the disk by the simplified model includes:

[0022] Calculating the blade root centrifugal tensile stress value of the blade by using the simplified model formula , wherein, is the blade root centrifugal tensile stress value of the blade, is the material density of the blade, is the blade outlet annular area, is the rotating speed of the gas turbine, is the shape factor of the blade;

[0023] Calculating the maximum shear stress of the disk by using the simplified model formula , wherein, is the maximum shear stress of the disk, is the Poisson's ratio, is the material density of the blade, is the rotating angular velocity of the blade, is the inner diameter of the disk, is the outer diameter of the disk, is the correction factor, which is related to the thickness variation curve of the disk, is the external load received by the outer edge of the disk, , is the number of turbine blades, is the centrifugal force of a single turbine blade acting on the disk, is the centrifugal force of a single rim boss acting on the disk;

[0024] Preliminary evaluation is made based on the blade root centrifugal tensile stress value of the blade and the maximum shear stress of the disk.

[0025] Further, the flow path layout, turbine diameter and blade height, blade axial width and axial spacing, meridian flow passage expansion angle are determined based on aerodynamic parameters and overall structural size constraints, including:

[0026] Designing the flow path layout based on the average radius height of the turbine, the flow area of the meridian flow passage and the structural length limit;

[0027] Designing the turbine diameter and blade height based on the aerodynamic performance and structural strength design requirements;

[0028] Designing the blade axial width based on the influence of turbine aspect ratio on turbine efficiency and the disk width limit caused by disk temperature and strength;

[0029] Designing the axial spacing of the blade based on the smooth transition of the flow path, the airflow excitation force, the effect of unsteady flow, the influence of potential effect and the restriction of the wheel disc structure;

[0030] Designing the expansion angle of the meridian flow passage based on the rationality of the blade outlet Mach number and the minimum expansion angle threshold;

[0031] The aerodynamic parameters include: flow coefficient, load coefficient, Mach number, reaction degree;

[0032] The overall structure size includes: the maximum outer diameter allowed by the cylinder, , the radius of the wheel disc, the width of the wheel disc, the axial length;

[0033] Wherein, is the product of the annular area at the outlet of the blade and the square of the rotational speed of the gas turbine.

[0034] Further, after completing the design of the meridian flow passage, the strength of the blade and the wheel disc is evaluated, including:

[0035] Based on the meridian flow passage size corresponding to the design result of the meridian flow passage, and by using the formula , the blade root centrifugal tensile stress value of the blade is calculated, wherein, is the updated shape factor of the blade, is the updated annular area at the outlet of the blade;

[0036] The blade root centrifugal tensile stress value is compared with the preset maximum allowable stress to determine whether the strength of the blade meets the requirements;

[0037] Based on the pre-established parameterized model of the wheel disc and the application of axisymmetric and plane stress assumption, the strength of the wheel disc is evaluated by the equal-thickness annular ring method.

[0038] The second aspect embodiment of the application proposes a heavy-duty gas turbine air-cooled turbine aerodynamic layout design system, including:

[0039] An acquisition module is configured to acquire the design requirements of a heavy-duty gas turbine air-cooled turbine, determine the design input parameters of the turbine based on the design requirements, and determine the number of turbine stages according to the design criteria, wherein the design input parameters include: inlet gas flow, total pressure, total temperature, power, efficiency, cooling gas flow and flow passage basic size;

[0040] A distribution module is configured to distribute the load of the turbine, determine the enthalpy drop and load distribution scheme of each turbine stage based on the cooling requirements of the downstream blade and the strength constraint of the blade;

[0041] The selecting module is used for one-dimensional design parameter selection, and is configured to select load coefficients, flow coefficients and reaction degrees of the turbine based on the load distribution result, and to calculate the product of the blade outlet annular area and the square of the gas turbine speed and the tip and rim speed, and to preliminarily evaluate the stress of the blade and the disc through a simplified model, and if the strength requirement is not met, the load coefficient and the flow coefficient are adjusted, and if the requirement is met, the first design module is entered;

[0042] The first design module is used for preliminary meridian flow passage design, and is configured to determine the flow path layout, the turbine diameter and the blade height, the blade axial width and the axial spacing and the meridian flow passage expansion angle based on the aerodynamic parameters and the overall structure size constraint, to evaluate the blade and disc strength after the meridian flow passage design is completed, to repeat the distribution module to the first design module for parameter optimization if the requirement is not met, and to enter the second design module if the requirement is met.

[0043] The second design module is used for flow pattern design, and is configured to determine the turbine stage reaction degree radial distribution law based on a controllable vortex design method for controlling the blade row outlet circulation.

[0044] The third design module is used for aerodynamic parameter radial distribution law design, and is configured to determine the radial distribution law of the aerodynamic parameters according to the one-dimensional average parameter design result.

[0045] The iteration module is used for S2 inverse problem calculation, and is configured to obtain the velocity triangle, the aerodynamic thermodynamic parameter and the detailed performance parameter of each radial section of the turbine stage through multiple iterations of the S2 flow surface.

[0046] The optimization module is used for aerodynamic performance analysis optimization, and is configured to optimize the average aerodynamic parameter and the radial distribution law thereof, to repeat the selecting module to the optimization module until the aerodynamic layout design result of the heavy-duty gas turbine gas-cooled turbine meeting the aerodynamic performance and strength requirements is obtained.

[0047] Preferably, the distribution module is further configured to:

[0048] Based on the downstream blade cooling requirement and the blade strength constraint, and using a step-by-step descending load distribution form, the enthalpy drop and the load distribution scheme of each stage are determined.

[0049] The third aspect of the embodiments of the present application provides an electronic device, which comprises a memory, a processor and a computer program stored in the memory and executable on the processor, and the processor implements the method according to the first aspect of the embodiments when executing the program.

[0050] The fourth aspect of the embodiments of the present application provides a computer readable storage medium having a computer program stored thereon, and the program is executed by a processor to implement the method according to the first aspect of the embodiments.

[0051] The technical scheme provided by the embodiments of the present application at least brings the following beneficial effects:

[0052] The application provides a heavy gas turbine air-cooled turbine aerodynamic layout design method, system, electronic equipment and storage medium. The method comprises the following steps: step 1: obtaining the design requirements of a heavy gas turbine air-cooled turbine, determining the design input parameters of the turbine based on the design requirements, and determining the turbine stage number according to the design criteria, wherein the design input parameters comprise inlet gas flow, total pressure, total temperature, power, efficiency, cold gas flow and flow channel basic size; step 2: turbine stage load distribution, determining the enthalpy drop and load distribution scheme of each turbine stage based on downstream blade cooling requirements and blade strength constraints; step 3: one-dimensional design parameter selection, selecting the load coefficient, flow coefficient and reaction degree of each stage of the turbine based on the stage load distribution result, calculating the product of the blade outlet annular area and the square of the turbine speed and the blade tip and rim speed, preliminarily evaluating the stress of the blade and the disc through a simplified model, adjusting the load coefficient and the flow coefficient if the strength requirement is not met, and entering step 4 if the requirement is met; step 4: initial meridian flow channel design, determining the flow path layout, turbine diameter and blade height, blade axial width and axial spacing and meridian flow channel expansion angle based on aerodynamic parameters and overall structural size constraints, evaluating the blade and disc strength after the meridian flow channel design is completed, repeating steps 2 to 4 for parameter optimization if the requirement is not met, and entering step 5 if the requirement is met; step 5: flow pattern design, determining the turbine stage reaction degree radial distribution law based on a controllable vortex design method for controlling the outlet annular quantity of the blade row; step 6: aerodynamic parameter radial distribution law design, determining the radial distribution law of the aerodynamic parameters according to the one-dimensional average parameter design result; step 7: S2 inverse problem calculation, obtaining the velocity triangle, aerodynamic thermodynamic parameters and detailed performance parameters of each radial section of the turbine stage through S2 flow surface multi-iteration calculation; step 8: aerodynamic performance analysis and optimization, optimizing the average aerodynamic parameters and the radial distribution law thereof, repeating steps 3 to 8 until the aerodynamic layout design result of the heavy gas turbine air-cooled turbine meeting the aerodynamic performance and strength requirements is obtained. The technical scheme provided by the application effectively eliminates the lag of aerodynamic design relative to the professional requirements of cooling, structure and strength, reduces the iteration number of turbine aerodynamic design, and shortens the turbine engineering design cycle.

[0053] Additional aspects and advantages of the application will be made apparent by the following description and the accompanying drawings. BRIEF DESCRIPTION OF DRAWINGS

[0054] The above and / or additional aspects and advantages of the application will become apparent and be readily understood from the following description, taken in conjunction with the accompanying drawings, in which:

[0055] Figure 1 A flow chart of a heavy gas turbine air-cooled turbine aerodynamic layout design method provided according to an embodiment of the application;

[0056] Figure 2 A turbine one-dimensional parameter design flow chart according to an embodiment of the present application is provided;

[0057] Figure 3 A turbine meridian flow passage design flow chart according to an embodiment of the present application is provided;

[0058] Figure 4 A detailed process flow chart of a heavy-duty gas turbine air-cooled turbine aerodynamic layout design method according to an embodiment of the present application is provided;

[0059] Figure 5 A structure diagram of a heavy-duty gas turbine air-cooled turbine aerodynamic layout design system according to an embodiment of the present application is provided. DETAILED DESCRIPTION

[0060] Embodiments of the present application are described in detail below with reference to examples illustrated in the attached drawings, in which the same or similar components have the same or similar designations throughout. The embodiments described below are examples and are intended to explain the present application, and are not to be understood as limiting the present application.

[0061] The heavy gas turbine air-cooled turbine aerodynamic layout design method, system, electronic device and storage medium provided by the application, the method comprises the following steps: step 1: obtaining the design requirements of the heavy gas turbine air-cooled turbine, and determining the design input parameters of the turbine based on the design requirements, and determining the turbine stage number according to the design criteria, wherein the design input parameters comprise: inlet gas flow, total pressure, total temperature, power, efficiency, cold gas flow and flow channel basic size; step 2: turbine stage load distribution, determining the enthalpy drop and load distribution scheme of each turbine stage based on the downstream blade cooling requirement and the blade strength constraint; step 3: one-dimensional design parameter selection, selecting the load coefficient, flow coefficient and reaction degree of each stage of the turbine based on the stage load distribution result, and calculating the product of the blade outlet annular area and the square of the gas turbine speed and the blade tip and rim speed, and preliminarily evaluating the stress of the blade and the disc through the simplified model, if the strength requirement is not met, the load coefficient and the flow coefficient are adjusted, if the requirement is met, step 4 is entered; step 4: initial meridian flow channel design, determining the flow path layout, turbine diameter and blade height, blade axial width and axial spacing, and meridian flow channel expansion angle based on the aerodynamic parameters and overall structure size constraint, evaluating the blade and disc strength after completing the meridian flow channel design, if the requirement is not met, repeating steps 2 to 4 to optimize the parameters, if the requirement is met, step 5 is entered; step 5: flow pattern design, determining the radial distribution law of the turbine stage reaction degree based on the controllable vortex design method of controlling the outlet annular quantity of the blade row; step 6: radial distribution law design of aerodynamic parameters, determining the radial distribution law of the aerodynamic parameters according to the one-dimensional average parameter design result; step 7: S2 inverse problem calculation, obtaining the velocity triangle, aerodynamic thermodynamic parameters and detailed performance parameters of each radial section of the turbine stage through multiple iterations of S2 flow surface calculation; step 8: aerodynamic performance analysis and optimization, optimizing the average aerodynamic parameters and their radial distribution law, repeating steps 3 to 8 until the aerodynamic layout design result of the heavy gas turbine air-cooled turbine meeting the aerodynamic performance and strength requirements is obtained. The technical scheme provided by the application effectively eliminates the lag of aerodynamic design relative to the professional requirements of cooling, structure and strength, reduces the number of iterations of turbine aerodynamic design, and shortens the turbine engineering design cycle.

[0062] The heavy gas turbine air-cooled turbine aerodynamic layout design method, system, electronic device and storage medium of the embodiment of the application are described below with reference to the accompanying drawings.

[0063] Embodiment one

[0064] Figure 1 A flow chart of a heavy gas turbine air-cooled turbine aerodynamic layout design method provided according to an embodiment of the application is shown in Figure 1 The method comprises the following steps:

[0065] Step 1: Obtain the design requirements of the heavy gas turbine air-cooled turbine, and determine the design input parameters of the turbine based on the design requirements, and determine the number of turbine stages according to the design criteria, wherein the design input parameters include: inlet gas flow, total pressure, total temperature, power, efficiency, cold gas flow and flow passage basic size;

[0066] Step 2: Load distribution of turbine stages, determine the enthalpy drop and load distribution scheme of each turbine stage based on the cooling requirements of downstream blades and the strength constraints of blades;

[0067] In the embodiments of the present disclosure, the determination of the enthalpy drop and load distribution scheme of each turbine stage based on the cooling requirements of downstream blades and the strength constraints of blades comprises:

[0068] Based on the cooling requirements of downstream blades and the strength constraints of blades, and using a step-by-step descending load distribution form, the enthalpy drop and load distribution scheme of each stage is determined.

[0069] It should be noted that based on the cooling requirements of downstream blades and the strength constraints of blades, the cold gas flow distribution of different sections and its influence on the temperature drop of the main gas by mixing with the gas are considered, and a step-by-step descending load distribution form is used to determine the enthalpy drop and load distribution scheme of each stage.

[0070] Among them, the inlet temperature and pressure of the first stage turbine are relatively high, and the working medium has strong power, so the power distribution of the first stage turbine is appropriately increased, thereby reducing the inlet pressure and temperature of the second stage turbine, so as to reduce the cooling design pressure of the second stage turbine.

[0071] It should be noted that the load distribution of the multi-stage turbine determines the distribution of the expansion ratio and the load size of each stage, and also determines the temperature drop distribution of the turbine stage, which will affect the cooling design boundary conditions of the downstream blades. Therefore, the cooling design requirements of the downstream blades are comprehensively considered, and the enthalpy drop and load distribution scheme with high aerodynamic efficiency is obtained under the condition of meeting the strength constraints of the blades.

[0072] Specifically, the load distribution of the turbine stage adopts a step-by-step descending form. The inlet temperature and pressure of the first stage are relatively high, and the working medium has strong power, so the power distribution is slightly more, which can also reduce the inlet pressure and temperature of the second stage turbine, and reduce the cooling design pressure of the second stage. At the same time, the variable working condition performance of the turbine is considered, and the power distribution of the last stage should not be too small to prevent the last stage turbine from entering the compressor working state.

[0073] At the same time, when the turbine aerodynamic performance is designed, the cold gas flow distribution of different sections and its influence on the temperature drop of the main gas by mixing with the gas are considered.

[0074] Step 3: One-dimensional design parameter selection, based on the stage load distribution result, selecting the load coefficient, flow coefficient and reaction degree of each stage of the turbine, and calculating the product of the blade outlet annular area and the square of the gas turbine speed and the tip and rim speed, preliminarily evaluating the stress of the blade and the disc through a simplified model, if the strength requirement is not met, adjusting the load coefficient and the flow coefficient, if it is met, entering step 4;

[0075] In the embodiments of the present disclosure, the selecting the load coefficient, flow coefficient and reaction degree of each stage of the turbine based on the stage load distribution result comprises:

[0076] Selecting the load coefficient of each stage of the turbine based on the turbine mean diameter size and the turbine stage efficiency;

[0077] Selecting the flow coefficient based on the turbine blade height and the cascade outlet Mach number limit;

[0078] Selecting the reaction degree based on the rotor outlet Mach number, axial force and cooling design boundary conditions.

[0079] Further, the preliminarily evaluating the stress of the blade and the disc through a simplified model comprises:

[0080] Calculating the blade root centrifugal tensile stress value of the blade by using a simplified model formula , wherein, is the blade root centrifugal tensile stress value of the blade, is the material density of the blade, is the blade outlet annular area, is the gas turbine speed, is the shape factor of the blade;

[0081] Calculating the maximum shear stress of the disc by using a simplified model formula , wherein, is the maximum shear stress of the disc, is the Poisson's ratio, is the material density of the blade, is the blade angular velocity, is the disc inner diameter, is the disc outer diameter, is a correction factor, which is related to the disc thickness variation curve, is the external load received by the disc outer edge, , is the number of turbine blades, is the centrifugal force of a single turbine blade acting on the disc, is the centrifugal force of a single rim boss acting on the disc;

[0082] Preliminarily evaluating based on the blade root centrifugal tensile stress value of the blade and the maximum shear stress of the disc.

[0083] It should be noted that the one-dimensional design parameter selection. According to the preliminary stage load distribution, the design parameters of the turbine are reasonably selected, such as load coefficient, flow coefficient, reaction degree, etc., to determine the speed triangle at the middle diameter. At the same time, considering the corresponding constraints of structure and cooling, the one-dimensional design parameters are optimized to obtain higher aerodynamic efficiency under the condition of meeting the blade and disc strength. The one-dimensional design parameter selection process is shown in Figure 2 , and the specific process is as follows:

[0084] (1) Load coefficient selection. After the power distribution of each stage is determined, the load coefficient determines the turbine mid-diameter. Selecting a larger load coefficient is conducive to reducing the blade height, thereby reducing the load of the disc, and is conducive to the structure and strength design, but a larger load coefficient represents a more severe change of internal airflow parameters of the turbine, which is prone to flow losses such as airflow separation, shock wave, secondary flow, leakage, etc., and the aerodynamic design difficulty increases, which brings challenges to ensuring high efficiency. Selecting a smaller load coefficient is conducive to reducing the Mach number level in the cascade, thereby improving the turbine stage efficiency, but the radial size of the turbine increases, which has the technical risk of not meeting the blade and disc strength, and needs to be considered comprehensively in the design.

[0085] (2) Flow coefficient selection. After the flow is determined, the flow coefficient determines the blade height (the difference between the inner and outer diameters of the flow passage). Selecting a smaller flow coefficient is conducive to high-efficiency design, but it may cause the blade tensile stress to increase, which may cause problems in strength. A larger flow coefficient is conducive to reducing the blade height, but it increases the proportion of the end area, and the secondary flow has a serious impact, the Mach number in the cascade increases, and the aerodynamic design difficulty increases. Therefore, when selecting the flow coefficient, the turbine blade strength and the restriction of the Mach number at the outlet of the cascade are considered comprehensively.

[0086] (3) Reaction degree selection. The reaction degree mainly affects the Mach number at the outlet of the turbine guide vane and the moving vane, as well as the axial force. At the same time, the reaction degree also affects the boundary conditions of the moving vane cooling design. Selecting a larger reaction degree means that the fluid expands severely in the turbine moving vane, which may increase the leakage loss of the moving vane tip, and also increase the axial force of the moving vane. Selecting a smaller reaction degree is prone to cause difficulties in the flow organization at the root of the moving vane passage, but it can reduce the relative total temperature at the inlet of the moving vane.

[0087] (4) The value and tip and rim speed calculation. The turbine mid-diameter and blade height determine the (the product of the blade outlet annular area and the square of the gas turbine speed), which determines the ideal tensile stress level of the turbine blade root, and reflects the influence of the turbine flow passage size on the blade strength and life. The tip and rim speed depends on the size of the turbine outer diameter and inner diameter, and determines the maximum centrifugal stress of the turbine disc. Too high tip and rim speed will cause the strength problem of the turbine disc to be prominent. After the turbine mid-diameter and blade height are preliminarily determined, the The values ​​and blade tip and rim tangential velocities were determined, and a preliminary assessment was made based on data and design experience from similar turbines. Whether the values ​​of the blade tip and rim tangential velocity are feasible.

[0088] (5) Preliminary assessment of blade and disk strength. Based on the rim diameter and blade height, a simplified model is used to conduct a preliminary assessment of the stress on the blade and disk. If the requirements are not met, the load factor and flow factor are adjusted until the strength design requirements are met.

[0089] In the simplified model, formula (1) is used to estimate the centrifugal tensile stress at the leaf root:

[0090] (1)

[0091] In the formula, The material density of the blade ( ); The annular area of ​​the blade exit ( N is the engine speed ( ); is the shape factor of the leaf, which is related to the ratio of the leaf to the root tip cross-sectional area.

[0092] Leaf root stress and turbine The value is directly proportional to the blade tip-root area ratio and inversely proportional to the blade tip-root area ratio. During the design process, the turbine blade tip-root area ratio is determined based on structural, cooling design, and manufacturing limitations. The value is within a reasonable range, thereby controlling the stress level of the blade.

[0093] The turbine's last-stage blade exit has a large annular area, at which point... The value is relatively large. In order to reduce the root stress level of the blade, a smaller blade tip-to-root area ratio is usually required. However, since the last-stage blades need to maintain the axial exhaust of the airflow, the airflow turning angle at the blade root is relatively small, which is not conducive to increasing the blade root cross-sectional area. At this time, selecting too small a blade tip cross-sectional thickness may result in a situation where the blade crown cannot be designed. Therefore, a trade-off balance is made among the design parameters in the design process to obtain the optimal design scheme.

[0094] The turbine disk was simplified to a conical disk for stress assessment, and the maximum tangential stress occurred on the inner diameter. Based on the formula for calculating the shear stress at the inner diameter of a wheel of uniform thickness, and after correction according to the geometric parameters of the conical wheel, formula (2) is used to quickly estimate the maximum shear stress of the wheel:

[0095] (2)

[0096] The centrifugal load generated by the external load on the outer edge of the wheel disc, i.e. the blade and the rim boss located on the outer edge of the wheel disc, is estimated by formula (3):

[0097] (3)

[0098] Step 4: meridional passage design, determining the flow path layout, turbine diameter and blade height, blade axial width and axial spacing, meridional passage expansion angle based on aerodynamic parameters and overall structural size constraints, evaluating blade and disc strength after completing the meridional passage design, if not meeting the requirements, repeating steps 2 to 4 for parameter optimization, if meeting, entering step 5;

[0099] In the embodiment of the present disclosure, the determination of the flow path layout, turbine diameter and blade height, blade axial width and axial spacing, and meridional passage expansion angle based on aerodynamic parameters and overall structural size constraints comprises:

[0100] Designing the flow path layout based on the average radius height of the turbine, the flow area of the meridional passage, and the structural length limit;

[0101] Designing the turbine diameter and blade height based on the aerodynamic performance and structural strength design requirements;

[0102] Designing the blade axial width based on the influence of the turbine aspect ratio on the turbine efficiency and the wheel disc width limit generated by the wheel disc temperature and strength;

[0103] Designing the blade axial spacing based on the influence of the smooth transition of the flow path, the airflow excitation force, the unsteady flow effect, the potential effect, and the wheel disc structure limit;

[0104] Designing the meridional passage expansion angle based on the reasonableness of the blade outlet Mach number and the minimum expansion angle threshold;

[0105] The aerodynamic parameters include: flow coefficient, load coefficient, Mach number, reaction degree;

[0106] The overall structural size includes: the maximum outer diameter allowed by the cylinder, , wheel disc radius, wheel disc width, axial length;

[0107] Wherein, is the product of the blade outlet annular area and the square of the gas turbine speed.

[0108] It should be noted that the load coefficient is used to determine the turbine diameter, which in turn affects the flow path form and turbine diameter;

[0109] The flow coefficient is used to determine the blade height;

[0110] The reaction degree can affect the turbine stator and rotor outlet Mach number, and in turn affect the meridional passage expansion angle.

[0111] The maximum outer diameter allowed by the cylinder is a meridional flow passage design constraint;

[0112] The role is that the blade root stress is proportional to the value, and constraints are put forward for turbine diameter and blade height;

[0113] The wheel disc radius and the wheel disc width can affect the blade axial width and the axial spacing;

[0114] The axial length can affect the flow passage layout.

[0115] In the embodiment of the present disclosure, the blade and the wheel disc strength are evaluated after the meridional flow passage design is completed, which comprises:

[0116] Based on the meridional flow passage size corresponding to the meridional flow passage design result, the blade root centrifugal tensile stress value is calculated by using the formula , wherein, is the updated shape factor of the blade, is the updated blade outlet annular area;

[0117] The blade root centrifugal tensile stress value is compared with the preset maximum allowable stress to determine whether the strength of the blade meets the requirements;

[0118] It should be noted that the maximum allowable stress is determined by the relationship curve of the metal wall temperature and the creep stress, and the metal wall temperature is based on the turbine temperature drop distribution and the cooling air distribution, and is obtained by solving the heat balance equation of the main flow gas and the cooling air. If the blade root centrifugal tensile stress value is lower than the maximum allowable stress value, it is determined to meet the requirements; otherwise, it is determined not to meet the requirements;

[0119] Based on the wheel disc parameterization model established in advance and the axisymmetric and plane stress assumption, the wheel disc strength is evaluated by using the equal-thickness annular ring method, wherein the wheel disc parameterization model can be a two-dimensional simplified profile model of the wheel disc, which is determined by the radius parameter, the disc axial width parameter and the fillet radius size. The simplified wheel disc profile shape only contains straight lines and circular arcs, which are defined by key point coordinates, and other points on the wheel disc curve are obtained by analytical geometry algorithm.

[0120] It should be noted that the meridional flow passage design comprehensively considers the aerodynamic parameters such as the flow coefficient, the load coefficient, the Mach number, the reaction degree, and the overall structural size constraints such as the maximum outer diameter allowed by the cylinder, the turbine , the wheel disc radius, the wheel disc width, the axial length, etc., to determine the overall layout of the flow path, the width of the moving and static blades, the spacing between the moving and static blades, the average radius and height of each row of blade rows, the meridional flow passage expansion angle, etc., as shown in Figure 3 , which comprises:

[0121] ​(1) Flow path form determination. The layout of the flow path is determined according to the average radius height of the turbine, the flow area of the meridian flow path, and the structural length limit. The full three-dimensional flow effect and the blade and disc strength are comprehensively considered, and a composite form flow path design close to the equal inner diameter form is adopted.

[0122] (2) Turbine diameter and blade height determination. The selection of the turbine radial dimension takes into account the design requirements of aerodynamic performance and structural strength.

[0123] (3) Blade axial width determination. The influence of the turbine aspect ratio on the turbine efficiency is considered, and the disc width limit caused by the disc temperature and strength is also considered.

[0124] (4) Blade axial spacing determination. The influences of smooth transition of the flow path, airflow excitation force, unsteady flow effect, and potential effect are comprehensively considered, and the disc cavity sealing design and the disc width limit caused by the disc temperature and strength are also considered.

[0125] (5) Flow path profile and expansion angle determination. The relative reasonable Mach number level of the outlet of each row of blades is ensured, and the secondary flow control and loss control in the cascade are considered. For a turbine with given inlet and outlet flow areas, the size of the meridian flow path expansion angle directly determines the axial length of the turbine. The design of a smaller meridian flow path expansion angle can effectively control the secondary flow and loss, but will increase the axial length of the turbine.

[0126] (6) The shape and size of the meridian flow path have an important influence on the power distribution of each stage of the multi-stage turbine. In order to ensure the rationality of the power distribution and aerodynamic parameters of each stage, the geometric size of the meridian flow path is appropriately adjusted.

[0127] (7) Blade and disc strength evaluation. After the meridian flow path design is completed, the blade and disc strength is further evaluated. Based on the meridian flow path size design result, the centrifugal stress at the blade root is estimated by formula (1), and compared with the maximum allowable stress. The maximum allowable stress is determined by the relationship curve of the metal wall temperature and the creep stress, and the metal wall temperature is obtained by solving the heat balance equation of the main flow gas and the cooling air based on the turbine temperature drop distribution and the cooling air distribution.

[0128] For disc strength calculation, a parameterized model of the disc is established, and the axisymmetric and plane stress assumption is applied, and the disc is quickly evaluated by the equal thickness ring method, so as to more accurately predict the influence of the parameter change of the meridian flow path on the disc stress. In the one-dimensional strength calculation and rapid evaluation, only the influence of the main aerodynamic load (centrifugal load from the blade and tenon and groove disc rim protrusion and its own weight) is considered.

[0129] If the strength of the blade or disc does not meet the requirements, steps 2 to 4 are repeated to optimize the enthalpy drop and cold air distribution of the turbine stage, the one-dimensional design parameters, and the meridional flow channel parameters.

[0130] Step 5: Flow pattern design, the controllable vortex design method based on controlling the outlet circulation of the blade row is used to determine the radial distribution of the reaction degree of the turbine stage.

[0131] It should be noted that the flow pattern design. The controllable vortex design method based on controlling the outlet circulation of the turbine is used to carry out the flow pattern design, and the radial distribution of the reaction degree of each turbine stage is obtained. The radial distribution of the reaction degree determines the radial and blade radial modeling twist rules, and the manufacturability of the blade is considered to determine the radial distribution rule of the reaction degree.

[0132] Step 6: Radial distribution rule design of aerodynamic parameters, the radial distribution rule of the aerodynamic parameters is determined according to the one-dimensional average parameter design result;

[0133] Step 7: S2 inverse problem calculation, the velocity triangle, aerodynamic thermodynamic parameters and detailed performance parameters of each radial section of the turbine stage are obtained through multiple iterations of S2 flow surface calculation;

[0134] It should be noted that the basic equation group including the complete radial balance equation is solved, and the calculation result is used to evaluate the advantages and disadvantages of the one-dimensional design, feed back the performance of the one-dimensional design scheme, analyze and modify the one-dimensional design scheme.

[0135] Step 8: Aerodynamic performance analysis and optimization, the average aerodynamic parameters and their radial distribution rules are optimized, steps 3 to 8 are repeated until the aerodynamic layout design result of the heavy-duty gas turbine gas-cooled turbine meeting the aerodynamic performance and strength requirements is obtained.

[0136] It should be noted that the design result is comprehensively judged through the through-flow distribution and result review. The average aerodynamic parameters and their radial distribution rules are further optimized to obtain better aerodynamic performance, such as reaction degree, Mach number, gas flow turning angle and enthalpy drop rule.

[0137] Steps 2 to 8 are repeated to further optimize the one-dimensional design parameters and meridional flow channel layout until the turbine aerodynamic layout design result meeting the aerodynamic performance and blade and disc strength design requirements is obtained, and the aerodynamic efficiency is the highest.

[0138] It should be noted that the detailed process of the heavy-duty gas turbine gas-cooled turbine aerodynamic layout design method proposed in the embodiment based on the above scheme is as shown in Figure 4 .

[0139] The design method provided in the embodiment introduces the constraints of strength and structure on aerodynamic design in the low-dimensional design stage, selects one-dimensional aerodynamic parameters and meridian flow passage aerodynamic layout, and can effectively reduce the iteration number, shorten the design period, and improve the aerodynamic design level of the turbine.

[0140] In summary, the aerodynamic layout design method of the heavy-duty gas turbine air-cooled turbine provided in the embodiment effectively eliminates the lag of the aerodynamic design relative to the professional requirements of cooling, structure, strength, and the like, reduces the iteration number of the turbine aerodynamic design, and shortens the turbine engineering design period.

[0141] Embodiment Two

[0142] Figure 5 A structural diagram of an aerodynamic layout design system of a heavy-duty gas turbine air-cooled turbine according to an embodiment of the present application is shown in FIG. 1. Figure 5 As shown in FIG. 1, the system includes:

[0143] The obtaining module 100 is configured to obtain the design requirements of the heavy-duty gas turbine air-cooled turbine, determine the design input parameters of the turbine based on the design requirements, and determine the number of turbine stages according to the design criteria, wherein the design input parameters include: inlet gas flow, total pressure, total temperature, power, efficiency, cold gas flow, and flow passage basic size.

[0144] The distribution module 200 is configured to distribute the load of the turbine, determine the enthalpy drop and load distribution scheme of each turbine stage based on the downstream blade cooling requirement and blade strength constraint, and determine the load distribution scheme of each turbine stage.

[0145] The selection module 300 is configured to select one-dimensional design parameters, select the load coefficient, flow coefficient, and reaction degree of each stage of the turbine based on the load distribution result of the stage, calculate the product of the blade outlet annular area and the square of the turbine speed and the tip speed and the rim speed, preliminarily evaluate the stress of the blade and the disc through a simplified model, adjust the load coefficient and the flow coefficient if the strength requirement is not met, and enter the first design module 400 if the requirement is met.

[0146] The first design module 400 is configured to preliminarily design the meridian flow passage, determine the flow path layout, turbine diameter, and blade height, blade axial width and axial spacing, and meridian flow passage expansion angle based on the aerodynamic parameters and overall structure size constraint, evaluate the strength of the blade and the disc after completing the meridian flow passage design, return to the distribution module to the first design module 400 for parameter optimization if the requirement is not met, and enter the second design module 500 if the requirement is met.

[0147] The second design module 500 is configured to design the flow pattern, determine the radial distribution law of the turbine stage reaction degree based on the controllable vortex design method for controlling the outlet annular quantity of the blade row.

[0148] The third design module 600 is configured to design a radial distribution rule of aerodynamic parameters, and determine a radial distribution rule of the aerodynamic parameters according to a one-dimensional average parameter design result;

[0149] The iteration module 700 is configured to perform S2 inverse problem calculation, and obtain velocity triangles, aerodynamic thermodynamic parameters and detailed performance parameters of each radial section of the turbine stage through multiple rounds of iteration calculation of the S2 flow surface;

[0150] The optimization module 800 is configured to perform aerodynamic performance analysis and optimization, and optimize the average aerodynamic parameters and the radial distribution rule thereof. The selection module 300 to the optimization module 800 are repeatedly selected until a design result of the aerodynamic layout of the heavy-duty gas turbine gas-cooled turbine meeting the aerodynamic performance and strength requirements is obtained.

[0151] In the embodiment of the present disclosure, the distribution module 200 is further configured to:

[0152] Based on the cooling requirements of the downstream blades and the blade strength constraints, and by adopting a load distribution form of step-by-step reduction, an enthalpy drop and load distribution scheme of each stage is determined.

[0153] In the embodiment of the present disclosure, the distribution module 200 is further configured to:

[0154] Based on the cooling requirements of the downstream blades and the blade strength constraints, and by adopting a load distribution form of step-by-step reduction, an enthalpy drop and load distribution scheme of each stage is determined.

[0155] In the embodiment of the present disclosure, the selection module 300 is further configured to:

[0156] The load coefficient of each stage of the turbine is selected based on the size of the turbine center and the turbine stage efficiency.

[0157] The flow coefficient is selected based on the turbine blade height and the outlet Mach number limit of the cascade.

[0158] The degree of reaction is selected based on the outlet Mach number of the moving blade, the axial force and the cooling design boundary condition.

[0159] In the embodiment of the present disclosure, the selection module 300 is further configured to:

[0160] The centrifugal tensile stress value of the blade root is calculated by using a simplified model formula , wherein, is the centrifugal tensile stress value of the blade root, is the material density of the blade, is the outlet annular area of the blade, is the rotating speed of the gas turbine, is the shape factor of the blade;

[0161] The centrifugal tensile stress value of the blade root is calculated by using a simplified model formula calculating the maximum shear stress of the disk, wherein is the maximum shear stress of the disk, is the Poisson ratio, is the material density of the blade, is the angular velocity of the blade, is the inner diameter of the disk, is the outer diameter of the disk, is the correction factor, which is related to the thickness variation curve of the disk, is the external load received by the outer edge of the disk, , is the number of turbine blades, is the centrifugal force of a single turbine blade acting on the disk, is the centrifugal force of a single rim boss acting on the disk;

[0162] performing preliminary evaluation based on the blade root centrifugal tensile stress value of the blade and the maximum shear stress of the disk.

[0163] In the embodiments of the present disclosure, the first design module 400 is further used for:

[0164] designing the flow path layout based on the average radius height of the turbine, the flow area of the meridional flow passage, and the structural length limit;

[0165] designing the turbine diameter and the blade height based on the design requirements of aerodynamic performance and structural strength;

[0166] designing the blade axial width based on the influence of the turbine aspect ratio on the turbine efficiency and the disk width limit caused by the disk temperature and strength;

[0167] designing the blade axial spacing based on the influence of the smooth transition of the flow path, the airflow excitation force, the unsteady flow effect, the potential effect, and the structural limit of the disk;

[0168] designing the meridional flow passage expansion angle based on the reasonableness of the blade outlet Mach number and the minimum expansion angle threshold;

[0169] The aerodynamic parameters include: flow coefficient, load coefficient, Mach number, reaction degree;

[0170] The overall structural size includes: the maximum outer diameter allowed by the cylinder, , the disk radius, the disk width, and the axial length;

[0171] wherein, is the product of the blade outlet annular area and the square of the gas turbine speed.

[0172] In the embodiments of the present disclosure, the first design module 400 is further used for

[0173] Based on the meridian flow channel size corresponding to the design result of the meridian flow channel, the meridian flow channel size is calculated by using the formula , wherein the centrifugal tensile stress value of the blade root is calculated, is the shape factor of the updated blade, is the updated blade outlet annular area;

[0174] The centrifugal tensile stress value of the blade root is compared with the preset maximum allowable stress to determine whether the strength of the blade meets the requirements;

[0175] Based on the pre-established parameterized model of the wheel disc and the axisymmetric and plane stress assumption, the strength of the wheel disc is evaluated by the equal-thickness annular ring method.

[0176] In summary, the aerodynamic layout design system of the heavy gas turbine air-cooled turbine provided in the embodiment effectively eliminates the lag of the aerodynamic design relative to the professional requirements of cooling, structure, strength, etc., reduces the iteration number of turbine aerodynamic design, and shortens the turbine engineering design cycle.

[0177] Embodiment three

[0178] In order to realize the above-mentioned embodiments, the present disclosure further provides an electronic device, comprising a memory, a processor and a computer program stored in the memory and executable on the processor, wherein the processor executes the program to realize the method of embodiment one.

[0179] Embodiment four

[0180] In order to realize the above-mentioned embodiments, the present disclosure further provides a computer readable storage medium having a computer program stored thereon, wherein the program is executed by a processor to realize the method of embodiment one.

[0181] In the description of the present specification, the description of the terms "one embodiment", "some embodiments", "example", "specific example" or "some examples" means that the specific features, structures, materials or characteristics described in conjunction with the embodiment or example are included in at least one embodiment or example of the present application. In the present specification, the illustrative description of the above terms does not necessarily refer to the same embodiment or example. Moreover, the specific features, structures, materials or characteristics described can be combined in any suitable manner in any one or more embodiments or examples. In addition, the skilled in the art can combine and combine the different embodiments or examples described in the present specification and the features of the different embodiments or examples without contradiction.

[0182] Any processes or methods described in the flow charts or otherwise described herein can be understood as representing modules, segments, or portions of code that include one or more executable instructions for implementing the specified logical functions or steps, and the preferred embodiments of the application include additional or fewer steps, in other orders, with other functionality, in implementations of these preferred embodiments of the application. Thus, any of the steps, options, aspects, components, etc. discussed herein can be included or deleted in other embodiments of the application, and yet still be deemed to fall within the scope of the present application.

[0183] Although the embodiments of the present application have been shown and described above, it should be understood by those ordinary skilled in the art that the above embodiments are exemplary and cannot be construed as limiting the present application, and those ordinary skilled in the art can make changes, modifications, replacements and variations to the above embodiments within the scope of the present application.

Claims

1. A method of aerodynamic layout design of a gas-cooled turbine of a heavy-duty gas turbine, characterized in that, The method comprises: Step 1: obtaining the design requirements of the heavy gas turbine air-cooled turbine, and determining the design input parameters of the turbine based on the design requirements, and determining the turbine stage number according to the design criteria, wherein the design input parameters comprise: inlet gas flow, total pressure, total temperature, power, efficiency, cold gas flow and flow channel basic size; Step 2: turbine stage load distribution, determining the enthalpy drop and load distribution scheme of each turbine stage based on the downstream blade cooling requirement and the blade strength constraint; Step 3: one-dimensional design parameter selection, selecting the load coefficient, flow coefficient and reaction degree of each stage of the turbine based on the stage load distribution result, and calculating the product of the blade outlet annular area and the square of the turbine speed and the blade tip and rim speed, and preliminarily evaluating the stress of the blade and the disc through a simplified model, if the strength requirement is not met, adjusting the load coefficient and the flow coefficient, if it is met, entering step 4; Step 4: initial meridian flow passage design, determining the flow path layout, turbine diameter and blade height, blade axial width and axial spacing, and meridian flow passage expansion angle based on the aerodynamic parameter and overall structure size constraint, evaluating the blade and disc strength after completing the meridian flow passage design, if the requirement is not met, repeating steps 2 to 4 for parameter optimization, if it is met, entering step 5; Step 5: flow pattern design, determining the turbine stage reaction degree radial distribution law based on the controllable vortex design method of controlling the outlet annular quantity of the blade row; Step 6: aerodynamic parameter radial distribution law design, determining the radial distribution law of the aerodynamic parameter according to the one-dimensional average parameter design result; Step 7: S2 inverse problem calculation, obtaining the velocity triangle, aerodynamic thermodynamic parameter and detailed performance parameter of each radial section of the turbine stage through S2 flow surface multi-iteration calculation; Step 8: aerodynamic performance analysis and optimization, optimizing the average aerodynamic parameter and its radial distribution law, repeating steps 3 to 8 until the aerodynamic layout design result of the heavy gas turbine air-cooled turbine meeting the aerodynamic performance and strength requirements is obtained.

2. The method of claim 1, wherein, The determination of the enthalpy drop and load distribution scheme of each turbine stage based on the downstream blade cooling requirement and the blade strength constraint comprises: Based on the downstream blade cooling requirement and the blade strength constraint, and using the step-by-step descending load distribution form, the enthalpy drop and load distribution scheme of each stage is determined.

3. The method of claim 2, wherein, The selection of the load coefficient, flow coefficient and reaction degree of each stage of the turbine based on the stage load distribution result comprises: Selecting the load coefficient of each stage of the turbine based on the turbine mean diameter and turbine stage efficiency; Selecting the flow coefficient based on the turbine blade height and blade cascade outlet Mach number limit; Selecting the reaction degree based on the blade outlet Mach number, axial force and cooling design boundary condition.

4. The method of claim 3, wherein, The preliminary evaluation of the stress of the blade and the disc through a simplified model comprises: calculating a blade root centrifugal tensile stress value for the blade using a simplified model formula wherein is a blade root centrifugal tensile stress value, is a material density of the blade, is a blade exit annular area, is a gas turbine rotational speed, is a shape factor of the blade; using a simplified model formula , calculating the maximum shear stress of the disk, wherein, is the maximum shear stress of the disk, is the Poisson's ratio, is the material density of the blade, is the angular velocity of the blade rotation, is the inner diameter of the disk, is the outer diameter of the disk, is a correction factor related to the thickness variation curve of the disk, is the external load received by the outer edge of the disk, , is the number of turbine blades, is the centrifugal force of a single turbine blade acting on the disk, is the centrifugal force of a single rim boss acting on the disk; Preliminary evaluation based on the blade root centrifugal tensile stress value and the maximum shear stress of the disc.

5. The method of claim 4, wherein, The determination of the flow path layout, turbine diameter and blade height, blade axial width and axial spacing, and meridian flow passage expansion angle based on the aerodynamic parameter and overall structure size constraint comprises: Designing the flow path layout based on the turbine average radius height, meridian flow passage flow area and structure length limit; Designing the turbine diameter and blade height based on the aerodynamic performance and structure strength design requirement; Designing the axial width of the blade based on the influence of the turbine aspect ratio on the turbine efficiency and the width limitation of the wheel disc caused by the temperature and strength of the wheel disc; Designing the axial spacing of the blade based on the influence of the smooth transition of the flow path, the excitation force of the airflow, the effect of unsteady flow, and the potential effect and the structure of the wheel disc; Designing the meridian flow passage expansion angle based on the reasonableness of the blade outlet Mach number and the minimum expansion angle threshold; The aerodynamic parameters include: flow coefficient, load coefficient, Mach number, reaction degree; The overall structure dimensions include: the maximum outer diameter allowed by the cylinder, , the wheel disc radius, the wheel disc width, the axial length; wherein is the product of the annular area of the blade exit and the square of the gas turbine rotational speed.

6. The method of claim 5, wherein, After the meridian flow passage is designed, the strength of the blade and the wheel disc is evaluated, including: based on the meridian flow passage size corresponding to the meridian flow passage design result, and the formula , the blade root centrifugal tensile stress value of the blade is calculated, wherein, is the shape factor of the updated blade, is the updated blade outlet annular area; Comparing the centrifugal tensile stress value of the blade root with the preset maximum allowable stress to determine whether the strength of the blade meets the requirements; Based on the pre-established parameterized model of the wheel disc and the axisymmetric and plane stress assumption, the strength of the wheel disc is evaluated by the equal-thickness ring method.

7. A heavy-duty gas turbine aero-thermodynamic layout design system, characterized in that, The system comprises: An acquisition module is configured to acquire design requirements of the air-cooled turbine of the heavy-duty gas turbine, determine design input parameters of the turbine based on the design requirements, and determine the number of stages of the turbine according to design criteria, wherein the design input parameters include: inlet gas flow, total pressure, total temperature, power, efficiency, cooling gas flow, and flow passage basic size; A distribution module is configured to distribute the load of the turbine, determine the enthalpy drop and load distribution scheme of each stage of the turbine based on the cooling requirement of the downstream blade and the strength constraint of the blade, and determine the load distribution scheme of each stage of the turbine based on the cooling requirement of the downstream blade and the strength constraint of the blade. A selection module is configured to select one-dimensional design parameters, select the load coefficient, flow coefficient, and reaction degree of each stage of the turbine based on the load distribution result of each stage, calculate the product of the blade outlet annular area and the square of the speed of the gas turbine and the tip speed and the rim speed, and preliminarily evaluate the stress of the blade and the wheel disc by a simplified model. If the strength requirement is not met, the load coefficient and the flow coefficient are adjusted. If the requirement is met, the first design module is entered. The first design module is configured to preliminarily design the meridian flow passage, determine the flow path layout, turbine diameter, and blade height, blade axial width and axial spacing, and meridian flow passage expansion angle based on the aerodynamic parameters and overall structure size constraint. After the meridian flow passage is designed, the strength of the blade and the wheel disc is evaluated. If the requirement is not met, the parameter optimization is repeated from the distribution module to the first design module. If the requirement is met, the second design module is entered. The second design module is configured to design the flow pattern. The controllable vortex design method is used to determine the radial distribution law of the reaction degree of each stage of the turbine. The third design module is configured to design the radial distribution law of the aerodynamic parameters. The radial distribution law of the aerodynamic parameters is determined according to the one-dimensional average parameter design result. An iteration module is configured to perform S2 inverse problem calculation. The velocity triangle, aerodynamic thermodynamic parameter, and detailed performance parameter of each radial section of each stage of the turbine are obtained through multiple iterations of the S2 flow surface. An optimization module is configured to perform aerodynamic performance analysis and optimization. The average aerodynamic parameter and its radial distribution law are optimized. The selection module is repeatedly selected to the optimization module until the aerodynamic layout design result of the air-cooled turbine of the heavy-duty gas turbine meeting the aerodynamic performance and strength requirements is obtained.

8. The system of claim 7, wherein, The distribution module is further configured to: Determine the enthalpy drop and load distribution scheme of each stage based on the cooling requirement of the downstream blade and the strength constraint of the blade and using the step-by-step descending load distribution form.

9. An electronic device, comprising: It comprises: Memory, a processor, and a computer program stored on the memory and loadable on the processor, the processor implementing the method as claimed in any one of claims 1-6 when executing the program.

10. A computer-readable storage medium having stored thereon a computer program, characterized in that, The program is executed by the processor to implement the method as claimed in any one of claims 1-6.

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