A turbine blade wall thickness forward design method and system

Abstract

The application provides a turbine blade wall thickness positive design method and system, the method comprises the following steps: step 1: obtaining the type of turbine blade, the leading edge radius of the turbine blade, the trailing edge radius, the internal cooling mode of the turbine blade and the initial wall thickness of the leading edge, the initial wall thickness of the median chord and the initial wall thickness of the trailing edge of the turbine blade; step 2: performing aerodynamic evaluation on the turbine blade, judging whether the type meets the design requirements, if yes, entering step 3, otherwise, updating the type of turbine blade and returning to step 1; step 3: performing respective performance evaluation on the turbine blade, and judging whether the evaluation results are all qualified, if yes, entering step 4, otherwise, adjusting the initial wall thickness of the turbine blade and returning to step 1; step 4: taking the initial wall thickness of the leading edge, the initial wall thickness of the median chord and the initial wall thickness of the trailing edge of the turbine blade as the design wall thickness. The technical scheme provided by the application effectively shortens the design cycle.

Description

TECHNICAL FIELD

[0001] The present application relates to the technical field of turbine blade design, in particular to a turbine blade wall thickness forward design method and system. BACKGROUND

[0002] The traditional turbine blade wall thickness determination is a serial design. The design steps are as follows: first, cooling design is performed; then, structural integrity analysis is performed; after the cooling and structural integrity analysis indicators are qualified, casting and machining are performed. This serial method does not fully consider the requirements of multidisciplinary design and manufacturing at the initial design stage, resulting in repeated design iterations. The traditional wall thickness design method determines the wall thickness at the cooling scheme design stage. The cooling analysis is performed to determine whether the wall thickness is reasonable. If the wall thickness is not reasonable, the cooling design is returned to determine the wall thickness again. At the structural integrity analysis stage, the wall thickness is evaluated to determine whether it meets the strength requirements. If the wall thickness does not meet the strength requirements, the cooling scheme design stage is returned to determine the wall thickness again. At the casting stage, the foundry determines the wall thickness for casting. The wall thickness is often found to be unreasonable during the casting process, resulting in unqualified castings, affecting the quality and yield of the blades. At this time, the cooling scheme design stage is returned to determine the wall thickness again. At the machining stage, the foundry machines the cast blades. During the drilling process, the wall thickness is found to be unreasonable, resulting in some holes that cannot be drilled through. The cooling scheme design stage is returned to determine the wall thickness again. For example, the leading edge wall thickness of some blades is relatively thick. The film holes arranged at this position usually have an angle, resulting in a relatively long film hole length, and the film hole is more likely to be unable to drill through. As can be seen, the traditional design method may result in the inability to machine the blades consistent with the design due to the unreasonable wall thickness set at the initial cooling design stage, resulting in blade scrap, the need to return to the cooling design stage to adjust the wall thickness, and the resulting time and economic costs are huge. Therefore, there is an urgent need to propose a scheme that can shorten the blade design cycle. SUMMARY

[0003] The present application provides a turbine blade wall thickness forward design method and system to at least solve the technical problem of long design cycle and large economic cost caused by repeated iterations during blade design.

[0004] The first aspect embodiment of the present application provides a turbine blade wall thickness forward design method, which comprises the following steps:

[0005] Step 1: obtaining the type of turbine blade, the leading edge radius and the trailing edge radius of the turbine blade, the internal cooling method of the turbine blade, and the initial wall thickness of the leading edge, the initial wall thickness of the median chord, and the initial wall thickness of the trailing edge of the turbine blade;

[0006] Step 2: aerodynamic evaluation is performed on the turbine blade according to the type of the turbine blade, the leading edge radius of the turbine blade, the trailing edge radius of the turbine blade, the internal cooling mode of the turbine blade, and the initial wall thickness of the leading edge, the initial wall thickness of the mid-chord, and the initial wall thickness of the trailing edge of the turbine blade, to determine whether the type of the turbine blade meets the design requirements, if yes, proceed to Step 3, otherwise, update the type of the turbine blade and return to Step 1;

[0007] Step 3: one-dimensional cooling evaluation, one-dimensional structural integrity evaluation, casting feasibility evaluation, and machining feasibility evaluation are performed on the turbine blade based on the initial wall thickness of the leading edge, the initial wall thickness of the mid-chord, and the initial wall thickness of the trailing edge of the turbine blade, to obtain one-dimensional cooling evaluation results, one-dimensional structural integrity evaluation results, casting feasibility evaluation results, and machining feasibility evaluation results of the turbine blade, and to determine whether the evaluation results are all qualified, if yes, proceed to Step 4, otherwise, adjust the initial wall thickness of the leading edge, the initial wall thickness of the mid-chord, and the initial wall thickness of the trailing edge of the turbine blade and return to Step 1;

[0008] Step 4: the initial wall thickness of the leading edge, the initial wall thickness of the mid-chord, and the initial wall thickness of the trailing edge of the turbine blade are taken as the design wall thickness of the leading edge, the design wall thickness of the mid-chord, and the design wall thickness of the trailing edge of the turbine blade.

[0009] Preferably, the aerodynamic evaluation performed on the turbine blade according to the type of the turbine blade, the leading edge radius of the turbine blade, the trailing edge radius of the turbine blade, the internal cooling mode of the turbine blade, and the initial wall thickness of the leading edge, the initial wall thickness of the mid-chord, and the initial wall thickness of the trailing edge of the turbine blade to determine whether the type of the turbine blade meets the design requirements comprises:

[0010] When the type of the turbine blade is a stationary blade and the internal cooling of the leading edge region adopts a bushing impingement cooling, it is determined whether the wall thickness of the impingement region of the leading edge region is uniform, if yes, it is determined that the leading edge radius and the initial wall thickness of the leading edge meet the design requirements;

[0011] When the type of the turbine blade is a stationary blade, it is determined whether the wall thickness of the impingement region of the mid-chord region is uniform, if yes, it is determined that the initial wall thickness of the mid-chord meets the design requirements;

[0012] When the type of the turbine blade is a stationary blade and the internal cooling of the trailing edge region adopts a middle split slot gas outlet cooling form or a back cut gas outlet cooling form, it is determined whether the wall thickness of the impingement region of the trailing edge region is uniform, if yes, it is determined that the trailing edge radius and the initial wall thickness of the trailing edge meet the design requirements;

[0013] When the turbine blade is a rotor blade and the leading edge region is internally cooled by cast impingement cooling, it is determined whether the ratio of impingement distance to impingement width of the leading edge impingement region is greater than or equal to a preset first threshold value and less than or equal to a preset second threshold value, and if so, it is determined that the initial wall thickness of the leading edge meets the design requirements.

[0014] When the turbine blade is a rotor blade, it is determined whether the initial wall thickness of the mid-chord is uniform, and if so, it is determined that the initial wall thickness of the mid-chord meets the design requirements.

[0015] When the turbine blade is a rotor blade, it is determined whether the initial wall thickness of the trailing edge is uniform, and if so, it is determined that the initial wall thickness of the trailing edge meets the design requirements.

[0016] Further, the initial wall thickness of the leading edge, the initial wall thickness of the mid-chord, and the initial wall thickness of the trailing edge of the turbine blade are used to perform one-dimensional cooling evaluation, one-dimensional structural integrity evaluation, casting feasibility evaluation, and machining feasibility evaluation on the turbine blade, including:

[0017] The calculation points of the leading edge, the calculation points of the mid-chord, and the calculation points of the trailing edge of the turbine blade are obtained, and the initial metal wall thickness of each calculation point is determined based on the initial wall thickness of the leading edge, the initial wall thickness of the mid-chord, and the initial wall thickness of the trailing edge of the turbine blade. Then, the outer wall temperature and the inner wall temperature of each calculation point are determined based on the initial metal wall thickness of each calculation point, and one-dimensional cooling evaluation is performed on the turbine blade based on the outer wall temperature and the inner wall temperature of each calculation point.

[0018] The stress level of each evaluation section of the turbine blade obtained by one-dimensional cooling evaluation is obtained, and one-dimensional structural integrity evaluation is performed on the turbine blade based on the stress level of each evaluation section.

[0019] The initial wall thickness of the leading edge, the initial wall thickness of the mid-chord, and the initial wall thickness of the trailing edge of the turbine blade are used to perform casting feasibility evaluation on the turbine blade.

[0020] The casting deviation and the angle of the turbine blade are obtained to perform machining feasibility evaluation on the turbine blade, wherein the angle includes the inclination angle of the film hole and the compound angle of the film hole center line and the gas flow direction.

[0021] Further, the calculation formula of the outer wall temperature of each calculation point is as follows:

[0022]

[0023] In the formula, Tout,i is the outer wall temperature of the i-th calculation point, Tgas,i is the gas temperature at the i-th calculation point, Tgas is the gas heat exchange coefficient, a heat flow density of the i-th calculation point;

[0024] The calculation formula of the inner wall temperature of each calculation point is as follows:

[0025]

[0026] In the formula, an inner wall temperature of the i-th calculation point, a cold air temperature at the i-th calculation point, a cold air heat exchange coefficient, wherein, , an initial metal wall thickness of the i-th calculation point, a thermal conductivity of the metal.

[0027] Further, the one-dimensional cooling evaluation of the turbine blade based on the outer wall temperature and the inner wall temperature of each calculation point comprises:

[0028] determining a temperature gradient of each calculation point based on the outer wall temperature and the inner wall temperature of each calculation point;

[0029] judging whether the temperature gradient of each calculation point is within a preset first threshold range, if yes, determining that the one-dimensional cooling performance of the turbine blade meets the requirement, otherwise, the one-dimensional cooling performance of the turbine blade does not meet the requirement.

[0030] Further, the stress level of each evaluation section of the turbine blade obtained through the one-dimensional cooling evaluation is obtained, and a one-dimensional structural integrity evaluation of the turbine blade is performed based on the stress level of each evaluation section, comprising:

[0031] selecting N evaluation sections along the blade height direction of the blade body of the turbine blade, and extracting structural parameters of each evaluation section through a three-dimensional modeling software;

[0032] importing the blade profile and thickness of the turbine blade obtained through the one-dimensional cooling evaluation, a preset aerodynamic pressure load boundary condition, and a preset displacement constraint condition into a finite element software, and calculating an aerodynamic bending moment of each evaluation section;

[0033] importing the structural parameters of each evaluation section and the aerodynamic bending moment of each evaluation section into a one-dimensional static strength calculation software, and calculating a blade body average tensile stress and a bending stress level of the turbine blade;

[0034] judging whether the blade body average tensile stress and the bending stress level meet a preset safety margin design requirement, if yes, determining that the one-dimensional structural integrity performance of the turbine blade meets the requirement, otherwise, the one-dimensional structural integrity performance of the turbine blade does not meet the requirement;

[0035] The structure parameters of the check section include: radial height, section area, section gravity center coordinates, moment of inertia information, body gravity center coordinates, volume, and check point coordinates.

[0036] Further, the casting feasibility evaluation of the turbine blade based on the initial wall thickness of the leading edge, the initial wall thickness of the mid-chord, and the initial wall thickness of the trailing edge of the turbine blade comprises:

[0037] According to the alloy type, casting size, and complexity of the turbine blade, a minimum wall thickness limit and a maximum wall thickness control range are set;

[0038] The distribution of the initial wall thickness of the leading edge, the initial wall thickness of the mid-chord, and the initial wall thickness of the trailing edge of the turbine blade is smoothed in the CAD software;

[0039] The shrinkage defects and shrinkage porosity defects of thick areas or end areas are predicted by using casting simulation software, and whether the initial wall thickness of the leading edge, the initial wall thickness of the mid-chord, and the initial wall thickness of the trailing edge of the turbine blade meet the casting feasibility is evaluated based on the predicted defect results.

[0040] Further, the machining feasibility evaluation of the turbine blade by obtaining the casting wall thickness deviation and the angle of the turbine blade comprises:

[0041] The predicted film hole length of the leading edge of the turbine blade is determined according to the inclination angle of the film hole, the composite angle of the film hole center line and the gas flow direction, and the initial wall thickness of the leading edge of the turbine blade;

[0042] The actual wall thickness of the leading edge of the turbine blade is determined according to the casting wall thickness deviation of the turbine blade and the initial wall thickness of the leading edge of the turbine blade;

[0043] The actual film hole length of the leading edge of the turbine blade is determined according to the inclination angle of the film hole, the composite angle of the film hole center line and the gas flow direction, and the actual wall thickness of the leading edge of the turbine blade;

[0044] It is judged whether the difference between the actual film hole length and the predicted film hole length is less than or equal to ten percent of the predicted film hole length, if yes, it is determined that the turbine blade meets the machining feasibility, otherwise the turbine blade does not meet the machining feasibility.

[0045] Further, the adjustment of the initial wall thickness of the leading edge, the initial wall thickness of the mid-chord, and the initial wall thickness of the trailing edge of the turbine blade comprises:

[0046] When the one-dimensional cooling performance of the turbine blade does not meet the requirements, if the temperature gradient of the calculation point is greater than a preset first gradient threshold, the wall thickness of the calculation point is increased, and if the temperature gradient of the calculation point is less than a preset second gradient threshold, the wall thickness of the calculation point is reduced;

[0047] when the one-dimensional structural integrity performance of the turbine blade does not meet the requirements, increasing the initial wall thickness of the leading edge, the initial wall thickness of the mid-chord, and the initial wall thickness of the trailing edge of the turbine blade;

[0048] when the turbine blade does not meet the casting feasibility, adjusting the wall thickness based on a preset wall thickness casting increase / decrease scheme;

[0049] when the turbine blade does not meet the machining feasibility, reducing the initial wall thickness of the leading edge of the turbine blade.

[0050] The second aspect embodiment of the application provides a turbine blade wall thickness forward design system, comprising:

[0051] The acquisition module is configured to acquire the type of the turbine blade, the leading edge radius of the turbine blade, the trailing edge radius of the turbine blade, the internal cooling mode of the turbine blade, and the initial wall thickness of the leading edge, the initial wall thickness of the mid-chord, and the initial wall thickness of the trailing edge of the turbine blade.

[0052] The first evaluation module is configured to perform aerodynamic evaluation on the turbine blade according to the type of the turbine blade, the leading edge radius of the turbine blade, the trailing edge radius of the turbine blade, the internal cooling mode of the turbine blade, and the initial wall thickness of the leading edge, the initial wall thickness of the mid-chord, and the initial wall thickness of the trailing edge of the turbine blade, to determine whether the type of the turbine blade meets the design requirements.

[0053] The second evaluation module is configured to perform one-dimensional cooling evaluation, one-dimensional structural integrity evaluation, casting feasibility evaluation, and machining feasibility evaluation on the turbine blade based on the initial wall thickness of the leading edge, the initial wall thickness of the mid-chord, and the initial wall thickness of the trailing edge of the turbine blade, to obtain one-dimensional cooling evaluation results, one-dimensional structural integrity evaluation results, casting feasibility evaluation results, and machining feasibility evaluation results of the turbine blade, and to determine whether the evaluation results are all qualified.

[0054] The determination module is configured to take the initial wall thickness of the leading edge, the initial wall thickness of the mid-chord, and the initial wall thickness of the trailing edge of the turbine blade as the design wall thickness of the leading edge, the design wall thickness of the mid-chord, and the design wall thickness of the trailing edge of the turbine blade.

[0055] The technical scheme provided by the embodiments of the application at least brings the following beneficial effects:

[0056] The application provides a turbine blade wall thickness positive design method and system, the method comprises the following steps: step 1: obtaining the type of the turbine blade, the leading edge radius of the turbine blade, the trailing edge radius of the turbine blade, the internal cooling mode of the turbine blade and the initial wall thickness of the leading edge, the initial wall thickness of the median chord and the initial wall thickness of the trailing edge of the turbine blade; step 2: performing aerodynamic evaluation on the turbine blade according to the type of the turbine blade, the leading edge radius of the turbine blade, the trailing edge radius of the turbine blade, the internal cooling mode of the turbine blade and the initial wall thickness of the leading edge, the initial wall thickness of the median chord and the initial wall thickness of the trailing edge of the turbine blade, determining whether the type of the turbine blade meets the design requirements, if yes, proceeding to step 3, otherwise, updating the type of the turbine blade and returning to step 1; step 3: performing one-dimensional cooling evaluation, one-dimensional structural integrity evaluation, casting feasibility evaluation and machining feasibility evaluation on the turbine blade based on the initial wall thickness of the leading edge, the initial wall thickness of the median chord and the initial wall thickness of the trailing edge of the turbine blade, obtaining one-dimensional cooling evaluation results, one-dimensional structural integrity evaluation results, casting feasibility evaluation results and machining feasibility evaluation results of the turbine blade, and determining whether the evaluation results are all qualified, if yes, proceeding to step 4, otherwise, adjusting the initial wall thickness of the leading edge, the initial wall thickness of the median chord and the initial wall thickness of the trailing edge of the turbine blade and returning to step 1; step 4: taking the initial wall thickness of the leading edge, the initial wall thickness of the median chord and the initial wall thickness of the trailing edge of the turbine blade as the design wall thickness of the leading edge, the design wall thickness of the median chord and the design wall thickness of the trailing edge of the turbine blade. The technical scheme provided by the application shortens the design and manufacturing cycle and reduces the design and manufacturing cost.

[0057] Additional aspects and advantages of the application will be made apparent by the following description and the accompanying drawings. BRIEF DESCRIPTION OF DRAWINGS

[0058] The above and / or additional aspects and advantages of the application will become apparent and be made clear to those skilled in the art from the following description and the accompanying drawings, in which:

[0059] Figure 1 A flowchart of a turbine blade wall thickness positive design method according to an embodiment of the application is provided;

[0060] Figure 2 A turbine blade leading edge, median chord and trailing edge schematic diagram according to an embodiment of the application is provided;

[0061] Figure 3 A back cut out gas schematic diagram according to an embodiment of the application is provided;

[0062] Figure 4 A middle split slot out gas schematic diagram according to an embodiment of the application is provided;

[0063] Figure 5 A schematic diagram of calculating the distance from a point to a leading edge stagnation point according to an embodiment of the present application;

[0064] Figure 6 A schematic diagram of calculating a point according to an embodiment of the present application;

[0065] Figure 7 A schematic diagram of a bmm according to an embodiment of the present application;

[0066] Figure 8 A schematic diagram of casting bias and punching direction according to an embodiment of the present application;

[0067] Figure 9 A schematic diagram of a tilt angle according to an embodiment of the present application;

[0068] Figure 10 A schematic diagram of a compound angle according to an embodiment of the present application;

[0069] Figure 11 A structure diagram of a system for forward design of a turbine blade wall thickness according to an embodiment of the present application;

[0070] Reference Signs

[0071] Leading edge profile radius 1, trailing edge profile radius 2, mid-chord 3, casting impact hole 4, cavity 5, bulkhead 6, leading edge stagnation point 7, calculation point 8, tip process hole 9. DETAILED DESCRIPTION

[0072] Embodiments of the present application are described in detail below with reference to the attached drawing figures, wherein the same or like reference numerals are used throughout the figures to refer to the same or like elements or elements with the same or similar function. The embodiments described below are exemplary and are intended to be illustrative of the present application and are not to be construed as limiting thereof.

[0073] The turbine blade wall thickness positive design method and system provided by the application, the method comprises the following steps: step 1: obtaining the type of the turbine blade, the leading edge radius of the turbine blade, the trailing edge radius of the turbine blade, the internal cooling mode of the turbine blade and the initial wall thickness of the leading edge, the initial wall thickness of the median chord and the initial wall thickness of the trailing edge of the turbine blade; step 2: performing aerodynamic evaluation on the turbine blade according to the type of the turbine blade, the leading edge radius of the turbine blade, the trailing edge radius of the turbine blade, the internal cooling mode of the turbine blade and the initial wall thickness of the leading edge, the initial wall thickness of the median chord and the initial wall thickness of the trailing edge of the turbine blade, judging whether the type of the turbine blade meets the design requirements, if yes, entering step 3, otherwise, updating the type of the turbine blade and returning to step 1; step 3: performing one-dimensional cooling evaluation, one-dimensional structural integrity evaluation, casting feasibility evaluation and machining feasibility evaluation on the turbine blade based on the initial wall thickness of the leading edge, the initial wall thickness of the median chord and the initial wall thickness of the trailing edge of the turbine blade, obtaining the one-dimensional cooling evaluation result, the one-dimensional structural integrity evaluation result, the casting feasibility evaluation result and the machining feasibility evaluation result of the turbine blade, and judging whether the evaluation results are all qualified, if yes, entering step 4, otherwise, adjusting the initial wall thickness of the leading edge, the initial wall thickness of the median chord and the initial wall thickness of the trailing edge of the turbine blade and returning to step 1; step 4: taking the initial wall thickness of the leading edge, the initial wall thickness of the median chord and the initial wall thickness of the trailing edge of the turbine blade as the design wall thickness of the leading edge, the design wall thickness of the median chord and the design wall thickness of the trailing edge of the turbine blade. The technical scheme provided by the application shortens the design and manufacturing cycle and reduces the design and manufacturing cost.

[0074] A turbine blade wall thickness positive design method and system of an embodiment of the application will be described below with reference to the accompanying drawings.

[0075] Embodiment one

[0076] Figure 1 A flowchart of a turbine blade wall thickness positive design method provided according to an embodiment of the application is shown in Figure 1 The method comprises the following steps:

[0077] Step 1: obtaining the type of the turbine blade, the leading edge radius of the turbine blade, the trailing edge radius of the turbine blade, the internal cooling mode of the turbine blade and the initial wall thickness of the leading edge, the initial wall thickness of the median chord and the initial wall thickness of the trailing edge of the turbine blade;

[0078] Step 2: performing aerodynamic evaluation on the turbine blade according to the type of the turbine blade, the leading edge radius of the turbine blade, the trailing edge radius of the turbine blade, the internal cooling mode of the turbine blade and the initial wall thickness of the leading edge, the initial wall thickness of the median chord and the initial wall thickness of the trailing edge of the turbine blade, judging whether the type of the turbine blade meets the design requirements, if yes, entering step 3, otherwise, updating the type of the turbine blade and returning to step 1;

[0079] In the embodiments of the present disclosure, the turbine blade is aerodynamically evaluated according to the type of the turbine blade, the leading edge radius of the turbine blade, the trailing edge radius, the internal cooling mode of the turbine blade, and the initial wall thickness of the leading edge, the initial wall thickness of the mid-chord, and the initial wall thickness of the trailing edge of the leading edge of the turbine blade, to determine whether the type of the turbine blade meets the design requirements, including:

[0080] When the type of the turbine blade is a stationary blade and the internal cooling of the leading edge region adopts a bushing impingement cooling, it is determined whether the wall thickness of the impingement region of the leading edge region is uniform, and if so, it is determined that the leading edge radius and the initial wall thickness of the leading edge meet the design requirements.

[0081] When the type of the turbine blade is a stationary blade, it is determined whether the wall thickness of the impingement region of the mid-chord region is uniform, and if so, it is determined that the initial wall thickness of the mid-chord meets the design requirements.

[0082] When the type of the turbine blade is a stationary blade and the internal cooling of the trailing edge region adopts a middle split slot outflow cooling form or a back cut outflow cooling form, it is determined whether the wall thickness of the impingement region of the trailing edge region is uniform, and if so, it is determined that the trailing edge radius and the initial wall thickness of the trailing edge meet the design requirements.

[0083] When the type of the turbine blade is a moving blade and the internal cooling of the leading edge region adopts a casting impingement cooling form, it is determined whether the ratio of the impingement distance to the impingement width of the leading edge impingement region is greater than or equal to a preset first threshold value and less than or equal to a preset second threshold value, and if so, it is determined that the initial wall thickness of the leading edge meets the design requirements.

[0084] It should be noted that, as shown in Figure 3 the impingement width is Lw and the impingement distance is Lj. To ensure the impingement cooling effect, the ratio of the impingement distance to the impingement width of the leading edge impingement region needs to be greater than or equal to a preset first threshold value and less than or equal to a preset second threshold value.

[0085] When the type of the turbine blade is a moving blade, it is determined whether the initial wall thickness of the mid-chord is uniform, and if so, it is determined that the initial wall thickness of the mid-chord meets the design requirements.

[0086] When the type of the turbine blade is a moving blade, it is determined whether the initial wall thickness of the trailing edge is uniform, and if so, it is determined that the initial wall thickness of the trailing edge meets the design requirements.

[0087] It should be noted that the first step of blade design is to propose an internal cooling structure arrangement scheme and wall thickness distribution according to the blade type, considering whether the blade type is a moving blade or a stationary blade, and what kind of cooling scheme is adopted for the leading edge, the mid-chord, and the trailing edge. This step determines whether the blade type meets the cooling design requirements according to the designed cooling scheme and wall thickness, and if not, the aerodynamic design of the blade type is updated.

[0088] For example, such as Figure 2 and Figure 3 As shown, in the leading edge region: For the stationary blade, the leading edge airfoil radius 1 is relatively large, and the entire blade body is usually cooled by bushing impact, resulting in a relatively uniform wall thickness distribution from the leading edge to the trailing edge. For the moving blade, the leading edge airfoil radius is usually smaller. If casting impact holes 4 are used for impact cooling, there are certain requirements for parameters such as the impact distance Lj and the radius of the impact target surface, which will result in a relatively thicker leading edge wall thickness compared to other parts of the blade body. When determining the cooling scheme, if the leading edge airfoil radius cannot meet the wall thickness requirements, resulting in an excessively thick or thin wall thickness, feedback is needed to modify the airfoil.

[0089] In the trailing edge region, whether it is a moving blade or a stationary blade, the cooling method (such as airflow through the central slit or back-cut airflow) should be considered based on the trailing edge blade radius and the cooling method. Figure 3 and Figure 4 (As shown) Determine whether the trailing edge wall thickness meets the layout requirements. If it does not, it indicates that the trailing edge radius does not meet the cooling design requirements, and feedback is needed to modify the aerodynamic blade profile.

[0090] in, Figure 3 The cooling method is back-cut exhaust. Figure 4 The cooling method is air outlet through a central slit;

[0091] It should be noted that the blade's interior includes: chamber 5 and septum 6.

[0092] Step 3: Based on the initial wall thickness of the leading edge and the initial wall thickness of the middle chord of the turbine blade (wherein, Figure 2 The initial wall thickness of the leading edge, middle edge, and trailing edge of the turbine blade (reference number 3) is used to perform one-dimensional cooling assessment, one-dimensional structural integrity assessment, casting feasibility assessment, and machining feasibility assessment on the turbine blade. The results of the one-dimensional cooling assessment, one-dimensional structural integrity assessment, casting feasibility assessment, and machining feasibility assessment of the turbine blade are obtained. It is then determined whether all the assessment results are qualified. If they are, proceed to step 4; otherwise, adjust the initial wall thickness of the leading edge, the initial wall thickness of the middle edge, and the initial wall thickness of the trailing edge of the turbine blade, and return to step 1.

[0093] In this embodiment of the disclosure, the one-dimensional cooling assessment, one-dimensional structural integrity assessment, casting feasibility assessment, and machining feasibility assessment of the turbine blade based on the initial wall thickness of the leading edge, the initial wall thickness of the middle chord, and the initial wall thickness of the trailing edge of the turbine blade include:

[0094] 1) obtaining each calculation point of the leading edge, the middle chord and the trailing edge of the turbine blade, determining the initial metal wall thickness of each calculation point based on the initial wall thickness of the leading edge, the initial wall thickness of the middle chord and the initial wall thickness of the trailing edge of the turbine blade, then determining the outer wall temperature and the inner wall temperature of each calculation point based on the initial metal wall thickness of each calculation point, and performing one-dimensional cooling evaluation on the turbine blade based on the outer wall temperature and the inner wall temperature of each calculation point;

[0095] In the embodiments of the present disclosure, the calculation formula of the outer wall temperature of each calculation point is as follows:

[0096]

[0097] In the formula, is the outer wall temperature of the i th calculation point, is the gas temperature at the i th calculation point, is the gas heat transfer coefficient, is the heat flux density of the i th calculation point;

[0098] The calculation formula of the inner wall temperature of each calculation point is as follows:

[0099]

[0100] In the formula, is the inner wall temperature of the i th calculation point, is the cold gas temperature at the i th calculation point, is the cold gas heat transfer coefficient, wherein, , is the initial metal wall thickness of the i th calculation point, is the thermal conductivity of the metal.

[0101] Further, the one-dimensional cooling evaluation on the turbine blade based on the outer wall temperature and the inner wall temperature of each calculation point comprises:

[0102] determining the temperature gradient of each calculation point based on the outer wall temperature and the inner wall temperature of each calculation point;

[0103] judging whether the temperature gradient of each calculation point is within a preset first threshold range, if yes, determining that the one-dimensional cooling performance of the turbine blade meets the requirements, otherwise, the one-dimensional cooling performance of the turbine blade does not meet the requirements.

[0104] Specifically, for one-dimensional cooling evaluation, it comprises:

[0105] a1. Calculate the recovery temperature and heat transfer coefficient of the gas along the blade profile according to the aerodynamic parameters (gas inlet temperature, pressure, Mach number, flow angle, outlet back pressure, etc.). The aerodynamic parameter distribution of the gas along the blade profile can be calculated using aerodynamic analysis tools (such as mises), i.e. the recovery temperature, velocity, density, and various physical parameters of the gas at any calculation point, such as thermal conductivity, Pr number, etc. The gas-side heat transfer coefficient Hgas at any calculation point can be obtained through the external heat transfer calculation formula. The calculation of the external heat transfer coefficient uses different calculation formulas at different positions, such as the cylindrical disturbance formula for the leading edge region , the pressure surface, the suction surface, and the trailing edge region can use the external flat plate formula , is the thermal conductivity of the gas, D is the diameter of the leading edge, is the distance from calculation point 8 to leading edge stagnation point 7, as shown in Figure 5 .

[0106] b1. Select appropriate cooling correlation formula to calculate the cooling gas-side heat transfer coefficient according to the amount of cooling gas provided by the secondary air system, the cooling gas temperature, and the internal cooling structure selection;

[0107] If it is forced convection cooling, the pipe flow formula can be used: , d is the equivalent diameter of the internal cooling structure;

[0108] If it is impingement cooling, the impingement cooling formula can be used: ;

[0109] It should be noted that , , , , , , , , , , , are coefficients, and the values of each coefficient are determined based on the Reynolds number, wherein the calculation formula of the Reynolds number Re is , ρ is the density of the gas, v is the gas velocity, μ is the gas viscosity coefficient, which are provided by aerodynamics, l is the characteristic length, which is the diameter of the leading edge when the cylindrical disturbance formula is used, and the flow direction distance from the blade profile to the calculation point when the external flat plate formula is used;

[0110] c1. Calculate the one-dimensional distribution of wall temperature at different positions according to the one-dimensional heat conduction formula.

[0111] The inner wall temperature and the outer wall temperature can be calculated through the heat balance principle.

[0112] First, calculate the heat flux density as: wherein is the thermal conductivity of the metal, is the metal wall thickness.

[0113] The outer wall temperature is calculated according to the heat flux density: The inner wall temperature is calculated according to the heat flux density: ;

[0114] The selection of the calculation points is shown in Figure 6 : for the leading edge passage, at least three points of the leading edge, the pressure surface and the suction surface are selected; for the mid-chord passage, at least one point of the pressure surface and one point of the suction surface are selected in each passage; for the trailing edge, at least one point of the pressure surface and one point of the suction surface are selected for the mid-split cooling scheme, and at least one point of the suction surface is selected for the back-cut cooling scheme.

[0115] The outer wall temperature and the inner wall temperature of each calculation point are calculated through one-dimensional evaluation, and the rationality of the wall temperature distribution is judged.

[0116] 2) obtain the stress level of each examination section of the turbine blade through one-dimensional cooling evaluation, and perform one-dimensional structural integrity evaluation on the turbine blade based on the stress level of each examination section;

[0117] In the embodiment of the present disclosure, the obtaining of the stress level of each examination section of the turbine blade through one-dimensional cooling evaluation, and the one-dimensional structural integrity evaluation on the turbine blade based on the stress level of each examination section, comprises:

[0118] N examination sections are selected along the blade height direction of the blade body of the turbine blade, and the structural parameters of each examination section are extracted through a three-dimensional modeling software;

[0119] The blade profile and thickness of the turbine blade obtained through one-dimensional cooling evaluation, the preset aerodynamic pressure load boundary condition, and the preset displacement constraint condition are imported into a finite element software, and the aerodynamic bending moment of each examination section is calculated;

[0120] The structural parameters of each examination section and the aerodynamic bending moment of each examination section are imported into a one-dimensional static strength calculation software, and the average tensile stress and bending stress level of the blade body of the turbine blade are calculated;

[0121] It is judged whether the average tensile stress and bending stress level of the blade body meet the preset safety margin design requirement, if yes, it is determined that the one-dimensional structural integrity performance of the turbine blade meets the requirement, otherwise the one-dimensional structural integrity performance of the turbine blade does not meet the requirement;

[0122] The structural parameters of the examination section include: radial height, cross-sectional area, cross-sectional center of gravity coordinates, moment of inertia information, body center of gravity coordinates, volume, and examination point coordinates.

[0123] Specifically, one-dimensional structural integrity assessment is performed on the wall thickness to assess whether the strength meets the requirements. The assessment method is as follows: one-dimensional structural integrity strength assessment is performed according to the one-dimensional calculation results of cooling:

[0124] a2. At least five examination sections are selected along the height direction of the blade body. Structural parameters are extracted from the examination sections by a three-dimensional modeling software, including radial height, sectional area, sectional gravity center coordinates, moment of inertia information, body gravity center coordinates, volume, examination point coordinates, and the like.

[0125] b2. The blade profile structural model obtained through the aerodynamic and cooling design is imported into a finite element software, and aerodynamic bending moments corresponding to the examination sections are calculated by loading aerodynamic pressure load boundary conditions (provided by an aerodynamic professional) and displacement constraints (applied according to structural constraint relationships).

[0126] c2. The parameters described in a2 and b2 are imported into a one-dimensional static strength calculation software to obtain average tensile stress and bending stress levels of each section of the blade body, and whether the design requirements of safety margin are met is assessed. If the design requirements of safety margin are met, the wall thickness can meet the strength requirements. If the design requirements of safety margin are not met, wall thickness optimization suggestions are proposed, and the above contents are repeated and iterated with aerodynamics and cooling until the requirements are met.

[0127] Selection of calculation points: examination sections: at least five examination sections are selected along the height direction of the blade body; examination points: 1) leading edge: at least two points are selected; 2) trailing edge: at least two points are selected; 3) suction surface: at least one point is selected. The average tensile stress and bending stress levels of each calculation point are calculated through one-dimensional static strength assessment, and whether they are within the material allowable stress limit range and meet the safety margin design requirements is assessed.

[0128] 3) performing a casting feasibility assessment on the turbine blade based on the initial wall thickness of the leading edge, the initial wall thickness of the mid-chord, and the initial wall thickness of the trailing edge of the turbine blade;

[0129] In the embodiments of the present disclosure, the casting feasibility assessment on the turbine blade based on the initial wall thickness of the leading edge, the initial wall thickness of the mid-chord, and the initial wall thickness of the trailing edge of the turbine blade includes:

[0130] According to the alloy type, casting size, and complexity of the turbine blade, a minimum wall thickness limit and a maximum wall thickness control range are set;

[0131] The distribution of the initial wall thickness of the leading edge, the initial wall thickness of the mid-chord, and the initial wall thickness of the trailing edge of the turbine blade is smoothed in the CAD software;

[0132] The shrinkage defects and shrinkage porosity defects of thick or end regions are predicted by using casting simulation software, and whether the initial wall thickness of the leading edge, the initial wall thickness of the mid-chord, and the initial wall thickness of the trailing edge of the turbine blade meet the casting feasibility is evaluated based on the predicted defect results.

[0133] Specifically, the rationality of the wall thickness is considered from the casting perspective.

[0134] Firstly, the structural characteristics of the turbine blade are analyzed in detail, including complexity, surface change and key positions, and the key points and difficulties of the wall thickness design are determined, and then the wall thickness design is performed according to three principles: uniformity principle, minimum wall thickness limit, and maximum wall thickness control.

[0135] Uniformity principle: under the premise of meeting the structural strength and cooling demand, the wall thickness is as uniform as possible to avoid sudden changes in wall thickness that are too large or too small.

[0136] Minimum wall thickness limit: according to the casting process and material properties, the minimum wall thickness limit is set to ensure that there are no pouring defects, cold separation and other defects in the casting process.

[0137] Maximum wall thickness control: control the maximum wall thickness to prevent casting defects such as shrinkage and shrinkage porosity caused by excessive wall thickness.

[0138] Then, the designed wall thickness distribution is simulated and analyzed by using casting simulation software to predict the flow, solidification and defect formation during the casting process. And at the same time, the accumulated casting experience and structural design experience are combined to complete the turbine blade wall thickness design.

[0139] Through the above evaluation method, the rationality of the wall thickness distribution of the turbine blade cooling structure from the casting perspective can be considered comprehensively, and the design iteration and casting defects caused by unreasonable wall thickness can be effectively avoided.

[0140] The main operations are to determine the specific cast high-temperature alloy grade used by the blade, to obtain the material parameters that have a greater impact on the wall thickness design, such as thermal conductivity, shrinkage rate, density, etc., by consulting the casting manual, material database, etc. of the casting alloy; from the preliminary cooling structure design scheme, the key dimensions and curvatures are determined by using CAD software to measure and determine the key dimensions and curvatures, the complex cooling structure is classified and parameterized modeling, the smooth transition of the blade shape and wall thickness is realized; based on experience, combined with material data such as thermal expansion coefficient, simple simulation analysis is carried out.

[0141] According to the body shrinkage of the alloy, combined with the size of the cooling channel, it is preliminarily judged which areas (usually thick areas or end areas) are prone to shrinkage porosity, providing a basis for subsequent adjustment of wall thickness. Determine the cross-sectional shape of the cooling channel (circular, square, rectangular, irregular, etc.) and the preliminary cross-sectional size (diameter or side length). According to the functional requirements (such as flow, heat exchange efficiency) and preliminary strength considerations, set an initial wall thickness. Mark out the complex or prone to problems areas such as sharp turns of the channel, sudden changes in cross-section, connection with the external structure of the blade, connection between thin and thick walls, etc. According to the experience value of similar successful designs, or according to the preliminary strength estimate, set an initial wall thickness value for different areas. For example, for internal serpentine channels, the initial wall thickness may be set to 1.5mm-3.0mm (the specific value needs to be judged according to the alloy and size level).

[0142] According to the alloy type, casting size, complexity, refer to the casting manual or internal standard to determine the minimum wall thickness that the alloy can successfully cast. For example, for nickel-based high-temperature alloy with good fluidity, the minimum wall thickness of small complex castings may be between 1.0mm-2.0mm. The wall thickness near the blade tip process hole is not less than bmm, such as the wall thickness near the blade tip process hole 9 shown in Figure 7

[0143] Examine the preliminary set wall thickness, especially at the junction of the cooling channel and the solid, and the reinforcing ribs. If the wall thickness at a certain place is significantly greater than the surrounding area (for example, it is 2-3 times or more than the surrounding channel wall thickness), it needs to be considered whether it can be reduced. Instead of simply increasing the wall thickness, it can be allowed to have a certain thickness by optimizing the structure shape (such as using gradual transition instead of sudden change) and strengthening the feeding design.

[0144] In possible cases, try to make the wall thickness of the cooling channel uniform. Uniform wall thickness helps to achieve uniform cooling effect, more importantly, it can make each part solidify at the same time, reducing thermal stress caused by different solidification sequences. In the CAD software, smooth the preliminary wall thickness distribution to ensure uniformity of wall thickness. For example, use "offset surface" and adjust "distance variation rate" or "transition" options to make the wall thickness change more gently along the surface normal, avoiding sharp corners or step-like wall thickness changes. Set an allowed wall thickness variation rate (for example, the wall thickness of adjacent two sections does not change by more than a certain percentage or absolute value). In CAD or special wall thickness checking software, check the wall thickness gradient and manually adjust the areas that exceed the standard.

[0145] 4) Obtain the casting deviation and angle of the turbine blade, and evaluate the machining feasibility of the turbine blade, wherein the angle includes: the inclination angle of the film hole, and the complex angle of the film hole center line and the gas flow direction.

[0146] ​In the embodiments of the present disclosure, the casting wall thickness deviation of the turbine blade and the machining feasibility evaluation of the turbine blade are obtained, and the machining feasibility evaluation of the turbine blade comprises:

[0147] The expected film hole length of the leading edge of the turbine blade is determined according to the inclination angle of the film hole, the compound angle of the film hole center line and the gas flow direction and the initial wall thickness of the leading edge of the turbine blade;

[0148] The actual wall thickness of the leading edge of the turbine blade is determined according to the casting wall thickness deviation of the turbine blade and the initial wall thickness of the leading edge of the turbine blade;

[0149] The actual film hole length of the leading edge of the turbine blade is determined according to the inclination angle of the film hole, the compound angle of the film hole center line and the gas flow direction and the actual wall thickness of the leading edge of the turbine blade;

[0150] It is judged whether the difference between the actual film hole length and the expected film hole length is less than or equal to 10% of the expected film hole length, if yes, it is determined that the turbine blade meets the machining feasibility, otherwise the turbine blade does not meet the machining feasibility.

[0151] Specifically, as shown in Figure 8 , the rationality of the wall thickness is considered from the perspective of machining, and the evaluation method is: according to the casting deviation, the position and angle of the blade to be punched, the feasibility of punching is evaluated, and the rationality of the wall thickness is judged. The feasibility of punching at the position with thick leading edge wall thickness needs to be evaluated in particular in the machining evaluation.

[0152] The angle between the film hole center line and the tangential direction of the blade section is called the inclination angle of the film hole as shown in Figure 9 , which is denoted by ; the angle between the film hole center line and the gas flow direction is called the compound angle as shown in Figure 10 , which is denoted by , and the wall thickness of the blade is , so the length of the film hole can be calculated as: According to the casting experience, the wall thickness deviation of the punching position can be obtained, and the actual casting wall thickness can be estimated, so the actual punching length can be calculated.

[0153] It should be noted that if the actual opening length exceeds more than 10% of the design punching length, the design value of the wall thickness needs to be reduced or the punching position, angle and other parameters need to be adjusted.

[0154] After the evaluation of the wall thickness of the above five professional directions meets the requirements, the details of the cooling design, the structural integrity analysis, the casting and the machining steps are performed.

[0155] In the embodiments of the present disclosure, the adjusting the initial wall thickness of the leading edge, the initial wall thickness of the mid-chord, and the initial wall thickness of the trailing edge of the turbine blade comprises:

[0156] When the one-dimensional cooling performance of the turbine blade does not meet the requirement, if the temperature gradient of the calculation point is greater than the preset first gradient threshold, the wall thickness of the calculation point is increased, and if the temperature gradient of the calculation point is less than the preset second gradient threshold, the wall thickness of the calculation point is reduced.

[0157] When the one-dimensional structural integrity performance of the turbine blade does not meet the requirement, the initial wall thickness of the leading edge, the initial wall thickness of the mid-chord, and the initial wall thickness of the trailing edge of the turbine blade are increased.

[0158] When the turbine blade does not meet the casting feasibility, the wall thickness is adjusted based on a preset wall thickness casting increase and decrease scheme.

[0159] It should be noted that the preset wall thickness casting increase and decrease scheme comprises:

[0160] The initial wall thickness is reviewed, especially at positions such as the connection between the cooling channel and the entity and the reinforcing rib. If the wall thickness at a certain position is obviously greater than that of the surrounding area (for example, is 2-3 times or more than that of the surrounding channel wall thickness), it is necessary to consider whether it can be reduced. The existence of a certain thickness area can be allowed by optimizing the structure shape (such as using a gradual transition instead of a sudden change) and strengthening the feeding design.

[0161] When the turbine blade does not meet the machining feasibility, the initial wall thickness of the leading edge of the turbine blade is reduced.

[0162] Step 4: Taking the initial wall thickness of the leading edge, the initial wall thickness of the mid-chord, and the initial wall thickness of the trailing edge of the turbine blade as the design wall thickness of the leading edge, the design wall thickness of the mid-chord, and the design wall thickness of the trailing edge of the turbine blade.

[0163] In summary, the turbine blade wall thickness forward design method provided in the embodiments effectively avoids repeated iterations caused by unreasonable initial wall thickness design in the blade design process, shortens the design and manufacturing cycle, and reduces the design and manufacturing cost.

[0164] Embodiment Two

[0165] Figure 11 The structure diagram of a turbine blade wall thickness forward design system provided according to an embodiment of the present application is shown in FIG. 1. Figure 11 As shown in FIG. 1, the system comprises:

[0166] The obtaining module 100 is configured to obtain the type of the turbine blade, the leading edge radius of the turbine blade, the trailing edge radius of the turbine blade, the internal cooling mode of the turbine blade, and the initial wall thickness of the leading edge, the initial wall thickness of the mid-chord, and the initial wall thickness of the trailing edge of the turbine blade.

[0167] The first evaluation module 200 is configured to perform aerodynamic evaluation on the turbine blade according to the type of the turbine blade, the leading edge radius of the turbine blade, the trailing edge radius of the turbine blade, the internal cooling mode of the turbine blade, and the initial wall thickness of the leading edge, the initial wall thickness of the mid-chord, and the initial wall thickness of the trailing edge of the turbine blade, and determine whether the type of the turbine blade meets the design requirement. If yes, the second evaluation module is entered; otherwise, the type of the turbine blade is updated, and the obtaining module 100 is returned.

[0168] The second evaluation module 300 is configured to perform one-dimensional cooling evaluation, one-dimensional structural integrity evaluation, casting feasibility evaluation, and machining feasibility evaluation on the turbine blade based on the initial wall thickness of the leading edge, the initial wall thickness of the mid-chord, and the initial wall thickness of the trailing edge of the turbine blade, respectively, to obtain one-dimensional cooling evaluation results, one-dimensional structural integrity evaluation results, casting feasibility evaluation results, and machining feasibility evaluation results of the turbine blade, and determine whether the evaluation results are all qualified. If yes, the determination module is entered; otherwise, the initial wall thickness of the leading edge, the initial wall thickness of the mid-chord, and the initial wall thickness of the trailing edge of the turbine blade are adjusted, and the obtaining module 100 is returned.

[0169] The determination module 400 is configured to take the initial wall thickness of the leading edge, the initial wall thickness of the mid-chord, and the initial wall thickness of the trailing edge of the turbine blade as the design wall thickness of the leading edge, the design wall thickness of the mid-chord, and the design wall thickness of the trailing edge of the turbine blade.

[0170] In the embodiments of the present disclosure, the first evaluation module 200 is further configured to:

[0171] When the type of the turbine blade is a stator blade and the internal cooling of the leading edge region adopts the bushing impingement cooling, it is determined whether the wall thickness of the impingement region of the leading edge region is uniform. If yes, it is determined that the leading edge radius and the initial wall thickness of the leading edge meet the design requirement.

[0172] When the type of the turbine blade is a stator blade, it is determined whether the wall thickness of the impingement region of the mid-chord region is uniform. If yes, it is determined that the initial wall thickness of the mid-chord meets the design requirement.

[0173] When the type of the turbine blade is a stator blade and the internal cooling of the trailing edge region adopts the intermediate split slot outflow cooling form or the back cut outflow cooling form, it is determined whether the wall thickness of the impingement region of the trailing edge region is uniform. If yes, it is determined that the trailing edge radius and the initial wall thickness of the trailing edge meet the design requirement.

[0174] When the type of the turbine blade is a rotor blade and the internal cooling of the leading edge region adopts the casting impingement cooling form, it is determined whether the ratio of the impingement distance to the impingement width of the leading edge impingement region is greater than or equal to a preset first threshold value and less than or equal to a preset second threshold value. If yes, it is determined that the initial wall thickness of the leading edge meets the design requirement.

[0175] When the type of the turbine blade is a moving blade, it is judged whether the initial wall thickness of the mid-chord is uniform, and if uniform, it is judged that the initial wall thickness of the mid-chord meets the design requirement;

[0176] When the type of the turbine blade is a moving blade, it is judged whether the initial wall thickness of the trailing edge is uniform, and if uniform, it is judged that the initial wall thickness of the trailing edge meets the design requirement.

[0177] In the embodiments of the present disclosure, the second evaluation module 300 is further configured to:

[0178] obtain each calculation point of the leading edge, each calculation point of the mid-chord and each calculation point of the trailing edge of the turbine blade, and determine the initial metal wall thickness of each calculation point based on the initial wall thickness of the leading edge, the initial wall thickness of the mid-chord and the initial wall thickness of the trailing edge of the turbine blade, then determine the outer wall temperature and the inner wall temperature of each calculation point based on the initial metal wall thickness of each calculation point, and perform one-dimensional cooling evaluation on the turbine blade based on the outer wall temperature and the inner wall temperature of each calculation point;

[0179] wherein, the calculation formula of the outer wall temperature of each calculation point is as follows:

[0180]

[0181] In the formula, is the outer wall temperature of the i th calculation point, is the gas temperature at the i th calculation point, is the gas heat exchange coefficient, is the heat flux density of the i th calculation point;

[0182] The calculation formula of the inner wall temperature of each calculation point is as follows:

[0183]

[0184] In the formula, is the inner wall temperature of the i th calculation point, is the cold gas temperature at the i th calculation point, is the cold gas heat exchange coefficient, wherein, , is the initial metal wall thickness of the i th calculation point, is the thermal conductivity of the metal.

[0185] obtain the stress level of each evaluation section of the turbine blade through one-dimensional cooling evaluation, and perform one-dimensional structural integrity evaluation on the turbine blade based on the stress level of each evaluation section;

[0186] casting feasibility evaluation is performed on the turbine blade based on an initial wall thickness of a leading edge of the turbine blade, an initial wall thickness of a mean chord, and an initial wall thickness of a trailing edge;

[0187] casting deviation and angle of the turbine blade are acquired, and a machining feasibility evaluation is performed on the turbine blade based on the casting deviation and the angle, wherein the angle includes a complex angle of a film hole center line and a gas flow direction.

[0188] Further, the second evaluation module 300 is further used for:

[0189] a temperature gradient of each calculation point is determined based on an outer wall temperature and an inner wall temperature of the calculation point;

[0190] whether the temperature gradient of each calculation point is within a preset first threshold range is determined, if yes, it is determined that a one-dimensional cooling performance of the turbine blade meets a requirement, otherwise, the one-dimensional cooling performance of the turbine blade does not meet the requirement.

[0191] Further, the second evaluation module 300 is further used for:

[0192] N examination sections are selected along a blade height direction of the turbine blade, and structural parameters of each examination section are extracted through a three-dimensional modeling software;

[0193] a blade profile and thickness of the turbine blade evaluated by the one-dimensional cooling evaluation, a preset aerodynamic pressure load boundary condition, and a preset displacement constraint condition are imported into a finite element software, and an aerodynamic bending moment of each examination section is calculated;

[0194] the structural parameters of each examination section and the aerodynamic bending moment of each examination section are imported into a one-dimensional static strength calculation software, and an average tensile stress and a bending stress level of the blade of the turbine blade are calculated;

[0195] whether the average tensile stress and the bending stress level of the blade meet a preset safety margin design requirement is determined, if yes, it is determined that a one-dimensional structural integrity performance of the turbine blade meets a requirement, otherwise, the one-dimensional structural integrity performance of the turbine blade does not meet the requirement;

[0196] the structural parameters of the examination section include a radial height, a sectional area, a sectional gravity center coordinate, inertia moment information, a body gravity center coordinate, a volume, and an examination point coordinate.

[0197] Further, the second evaluation module 300 is further used for:

[0198] a minimum wall thickness limit and a maximum wall thickness control range are set according to an alloy type, a casting size, and a complexity of the turbine blade;

[0199] smooth the distribution of the initial wall thickness of the leading edge, the initial wall thickness of the mid-chord, and the initial wall thickness of the trailing edge of the turbine blade in the CAD software;

[0200] predict shrinkage defects and shrinkage porosity defects of the thick area or the end area by using casting simulation software, and evaluate whether the initial wall thickness of the leading edge, the initial wall thickness of the mid-chord, and the initial wall thickness of the trailing edge of the turbine blade meet the casting feasibility based on the predicted defect results.

[0201] Further, the second evaluation module 300 is further used for:

[0202] determining the predicted film hole length of the leading edge according to the inclination angle of the film hole, the compound angle of the film hole center line and the gas flow direction, and the initial wall thickness of the leading edge of the turbine blade;

[0203] determining the actual wall thickness of the leading edge according to the casting wall thickness deviation of the turbine blade and the initial wall thickness of the leading edge of the turbine blade;

[0204] determining the actual film hole length of the leading edge according to the inclination angle of the film hole, the compound angle of the film hole center line and the gas flow direction, and the actual wall thickness of the leading edge of the turbine blade;

[0205] judging whether the difference between the actual film hole length and the predicted film hole length is less than or equal to ten percent of the predicted film hole length, if yes, determining that the turbine blade meets the machining feasibility, otherwise, the turbine blade does not meet the machining feasibility.

[0206] Further, the second evaluation module 300 is further used for:

[0207] when the one-dimensional cooling performance of the turbine blade does not meet the requirement, if the temperature gradient of the calculation point is greater than a preset first gradient threshold, increasing the wall thickness of the calculation point, and if the temperature gradient of the calculation point is less than a preset second gradient threshold, reducing the wall thickness of the calculation point;

[0208] when the one-dimensional structural integrity performance of the turbine blade does not meet the requirement, increasing the initial wall thickness of the leading edge, the initial wall thickness of the mid-chord, and the initial wall thickness of the trailing edge of the turbine blade;

[0209] when the turbine blade does not meet the casting feasibility, adjusting the wall thickness based on a preset wall thickness casting increase and decrease scheme;

[0210] when the turbine blade does not meet the machining feasibility, reducing the initial wall thickness of the leading edge of the turbine blade.

[0211] In summary, the turbine blade wall thickness forward design system proposed in the embodiment shortens the design and manufacturing cycle and reduces the design and manufacturing cost.

[0212] In the description of the application, the description of the terms "one embodiment", "some embodiments", "an example", "a specific example", or "some examples" etc. means that the specific features, structures, materials or characteristics described in connection with the embodiment or example are included in at least one embodiment or example of the application. The illustrative description of the above terms in the specification does not necessarily refer to the same embodiment or example. Moreover, the specific features, structures, materials or characteristics described can be combined in any appropriate manner in any one or more embodiments or examples. In addition, different embodiments or examples described in the specification and the features of different embodiments or examples can be combined and combined by those skilled in the art without contradiction, if any.

[0213] Any process or method descriptions or descriptions of the flow diagrams in the specification or otherwise described herein can be understood as representing the steps of a method or process that can be implemented in code, which can be executed by a computer or processor, and which includes one or more steps for implementing the functionality or process as claimed by the application. The scope of preferred embodiments of the application encompasses not only the described embodiments but also equivalent methods and processes that can be implemented in hardware, software, or a combination of hardware and software.

[0214] Although the embodiments of the application have been shown and described above, it should be understood that the above-described embodiments are exemplary and are not to be construed as limiting the application, and those skilled in the art can make changes, modifications, replacements and variations to the above-described embodiments within the scope of the application.

Claims

1. A method of forward design of a turbine blade wall thickness, characterized by, The method comprises: Step 1: obtaining the type of the turbine blade, the leading edge radius of the turbine blade, the trailing edge radius of the turbine blade, the internal cooling mode of the turbine blade, and the initial wall thickness of the leading edge, the initial wall thickness of the mid-chord, and the initial wall thickness of the trailing edge of the turbine blade; Step 2: performing aerodynamic evaluation on the turbine blade according to the type of the turbine blade, the leading edge radius of the turbine blade, the trailing edge radius of the turbine blade, the internal cooling mode of the turbine blade, and the initial wall thickness of the leading edge, the initial wall thickness of the mid-chord, and the initial wall thickness of the trailing edge of the turbine blade, and determining whether the type of the turbine blade meets the design requirements; if yes, proceeding to Step 3, otherwise, updating the type of the turbine blade and returning to Step 1; Step 3: performing one-dimensional cooling evaluation, one-dimensional structural integrity evaluation, casting feasibility evaluation, and machining feasibility evaluation on the turbine blade based on the initial wall thickness of the leading edge, the initial wall thickness of the mid-chord, and the initial wall thickness of the trailing edge of the turbine blade, obtaining the one-dimensional cooling evaluation result, the one-dimensional structural integrity evaluation result, the casting feasibility evaluation result, and the machining feasibility evaluation result of the turbine blade, and determining whether the evaluation results are all qualified; if yes, proceeding to Step 4, otherwise, adjusting the initial wall thickness of the leading edge, the initial wall thickness of the mid-chord, and the initial wall thickness of the trailing edge of the turbine blade and returning to Step 1; Step 4: taking the initial wall thickness of the leading edge, the initial wall thickness of the mid-chord, and the initial wall thickness of the trailing edge of the turbine blade as the design wall thickness of the leading edge, the design wall thickness of the mid-chord, and the design wall thickness of the trailing edge of the turbine blade; wherein the aerodynamic evaluation on the turbine blade according to the type of the turbine blade, the leading edge radius of the turbine blade, the trailing edge radius of the turbine blade, the internal cooling mode of the turbine blade, and the initial wall thickness of the leading edge, the initial wall thickness of the mid-chord, and the initial wall thickness of the trailing edge of the turbine blade, and determining whether the type of the turbine blade meets the design requirements, comprises: when the type of the turbine blade is a stationary blade and the internal cooling of the leading edge region adopts a bushing impingement cooling, determining whether the wall thickness of the impingement region of the leading edge region is uniform; if yes, determining that the leading edge radius and the initial wall thickness of the leading edge meet the design requirements; when the type of the turbine blade is a stationary blade, determining whether the wall thickness of the impingement region of the mid-chord region is uniform; if yes, determining that the initial wall thickness of the mid-chord meets the design requirements; when the type of the turbine blade is a stationary blade and the internal cooling of the trailing edge region adopts a middle split slot outflow cooling form or a back cut outflow cooling form, determining whether the wall thickness of the impingement region of the trailing edge region is uniform; if yes, determining that the trailing edge radius and the initial wall thickness of the trailing edge meet the design requirements; when the type of the turbine blade is a moving blade and the internal cooling of the leading edge region adopts a casting impingement cooling form, determining whether the ratio of the impingement distance to the impingement width of the leading edge impingement region is greater than or equal to a preset first threshold value and less than or equal to a preset second threshold value; if yes, determining that the initial wall thickness of the leading edge meets the design requirements; when the type of the turbine blade is a moving blade, determining whether the initial wall thickness of the mid-chord meets the design requirements; if yes, determining that the initial wall thickness of the mid-chord meets the design requirements. When the turbine blade is a moving blade, it is determined whether the initial wall thickness of the trailing edge is uniform, and if so, the initial wall thickness of the trailing edge meets the design requirements.

2. The method of claim 1, wherein, The one-dimensional cooling evaluation, one-dimensional structural integrity evaluation, casting feasibility evaluation and machining feasibility evaluation of the turbine blade based on the initial wall thickness of the leading edge, the initial wall thickness of the mid-chord and the initial wall thickness of the trailing edge include: Obtaining each calculation point of the leading edge, each calculation point of the mid-chord and each calculation point of the trailing edge of the turbine blade, and determining the initial metal wall thickness of each calculation point based on the initial wall thickness of the leading edge, the initial wall thickness of the mid-chord and the initial wall thickness of the trailing edge of the turbine blade, and then determining the outer wall temperature and the inner wall temperature of each calculation point based on the initial metal wall thickness of each calculation point, and performing one-dimensional cooling evaluation on the turbine blade based on the outer wall temperature and the inner wall temperature of each calculation point; Obtaining the stress level of each examination section of the turbine blade through one-dimensional cooling evaluation, and performing one-dimensional structural integrity evaluation on the turbine blade based on the stress level of each examination section; Performing casting feasibility evaluation on the turbine blade based on the initial wall thickness of the leading edge, the initial wall thickness of the mid-chord and the initial wall thickness of the trailing edge of the turbine blade; Obtaining the casting deviation and the angle of the turbine blade to perform machining feasibility evaluation on the turbine blade, wherein the angle includes the inclination angle of the film hole and the complex angle of the film hole center line and the gas flow direction.

3. The method of claim 2, wherein, The calculation formula of the outer wall temperature of each calculation point is as follows: wherein Twiis the outer wall temperature at the i-th calculation point, Tgiis the gas temperature at the i-th calculation point, Kgiis the gas heat exchange coefficient, Jiis the heat flux density at the i-th calculation point; The calculation formula of the inner wall temperature of each calculation point is as follows: wherein Twi is the inner wall temperature at the i-th calculation point, Tci is the cold gas temperature at the i-th calculation point, Kci is the cold gas heat exchange coefficient, wherein, , Tw0is the initial metal wall thickness at the i-th calculation point, Km is the thermal conductivity of the metal.

4. The method of claim 3, wherein, The one-dimensional cooling evaluation of the turbine blade based on the outer wall temperature and the inner wall temperature of each calculation point includes: Determining the temperature gradient of each calculation point based on the outer wall temperature and the inner wall temperature of each calculation point; Determining whether the temperature gradient of each calculation point is within a preset first threshold range, and if so, determining that the one-dimensional cooling performance of the turbine blade meets the requirements, otherwise the one-dimensional cooling performance of the turbine blade does not meet the requirements.

5. The method of claim 4, wherein, The one-dimensional structural integrity evaluation of the turbine blade based on the stress level of each examination section of the turbine blade through one-dimensional cooling evaluation includes: Selecting N examination sections along the blade height direction of the turbine blade, and extracting the structural parameters of each examination section through a three-dimensional modeling software; Importing the blade shape and thickness of the turbine blade through one-dimensional cooling evaluation, a preset aerodynamic pressure load boundary condition and a preset displacement constraint condition into a finite element software to calculate the aerodynamic bending moment of each examination section; Importing the structural parameters of each examination section and the aerodynamic bending moment of each examination section into a one-dimensional static strength calculation software to calculate the average tensile stress and bending stress level of the blade of the turbine blade; Determining whether the average tensile stress and bending stress level of the blade meet the preset safety margin design requirements, and if so, determining that the one-dimensional structural integrity performance of the turbine blade meets the requirements, otherwise the one-dimensional structural integrity performance of the turbine blade does not meet the requirements; The structure parameters of the check section include: radial height, section area, section gravity center coordinates, moment of inertia information, body gravity center coordinates, volume, and check point coordinates.

6. The method of claim 5, wherein, The casting feasibility evaluation of the turbine blade based on the initial wall thickness of the leading edge, the initial wall thickness of the mid-chord, and the initial wall thickness of the trailing edge of the turbine blade includes: According to the alloy type, casting size, and complexity of the turbine blade, the minimum wall thickness limit and the maximum wall thickness control range are set; The distribution of the initial wall thickness of the leading edge, the initial wall thickness of the mid-chord, and the initial wall thickness of the trailing edge of the turbine blade is smoothed in the CAD software; The shrinkage defects and shrinkage porosity defects of thick areas or end areas are predicted by using a casting simulation software, and whether the initial wall thickness of the leading edge, the initial wall thickness of the mid-chord, and the initial wall thickness of the trailing edge of the turbine blade meet the casting feasibility is evaluated based on the predicted defect results.

7. The method of claim 6, wherein, The machining feasibility evaluation of the turbine blade based on the casting wall thickness deviation and the angle of the turbine blade includes: The predicted film hole length of the leading edge of the turbine blade is determined according to the inclination angle of the film hole, the composite angle of the film hole center line and the gas flow direction, and the initial wall thickness of the leading edge of the turbine blade; The actual wall thickness of the leading edge of the turbine blade is determined according to the casting wall thickness deviation and the initial wall thickness of the leading edge of the turbine blade; The actual film hole length of the leading edge of the turbine blade is determined according to the inclination angle of the film hole, the composite angle of the film hole center line and the gas flow direction, and the actual wall thickness of the leading edge of the turbine blade; It is judged whether the difference between the actual film hole length and the predicted film hole length is less than or equal to ten percent of the predicted film hole length, if yes, it is determined that the turbine blade meets the machining feasibility, otherwise the turbine blade does not meet the machining feasibility.

8. The method of claim 7, wherein, The adjustment of the initial wall thickness of the leading edge, the initial wall thickness of the mid-chord, and the initial wall thickness of the trailing edge of the turbine blade includes: When the one-dimensional cooling performance of the turbine blade does not meet the requirements, if the temperature gradient of the calculation point is greater than a preset first gradient threshold, the wall thickness of the calculation point is increased, and if the temperature gradient of the calculation point is less than a preset second gradient threshold, the wall thickness of the calculation point is reduced; When the one-dimensional structural integrity performance of the turbine blade does not meet the requirements, the initial wall thickness of the leading edge, the initial wall thickness of the mid-chord, and the initial wall thickness of the trailing edge of the turbine blade are increased; When the turbine blade does not meet the casting feasibility, the wall thickness is adjusted based on a preset wall thickness casting increase and decrease scheme; When the turbine blade does not meet the machining feasibility, the initial wall thickness of the leading edge of the turbine blade is reduced.

9. A turbine blade wall thickness forward design system based on the turbine blade wall thickness forward design method according to any one of claims 1 to 8, characterized by, The system includes: An acquisition module is configured to acquire the type of turbine blade, the radius of the leading edge and the trailing edge of the turbine blade, the internal cooling mode of the turbine blade, and the initial wall thickness of the leading edge, the initial wall thickness of the mid-chord, and the initial wall thickness of the trailing edge of the leading edge of the turbine blade; The first evaluation module is configured to perform aerodynamic evaluation on the turbine blade according to the type of the turbine blade, the leading edge radius of the turbine blade, the trailing edge radius of the turbine blade, the internal cooling mode of the turbine blade, and the initial wall thickness of the leading edge, the initial wall thickness of the mid-chord, and the initial wall thickness of the trailing edge of the turbine blade, to determine whether the type of the turbine blade meets the design requirement, and if yes, to enter the second evaluation module, otherwise, to update the type of the turbine blade and return to the acquisition module. The second evaluation module is configured to perform one-dimensional cooling evaluation, one-dimensional structural integrity evaluation, casting feasibility evaluation, and machining feasibility evaluation on the turbine blade based on the initial wall thickness of the leading edge, the initial wall thickness of the mid-chord, and the initial wall thickness of the trailing edge of the turbine blade, to obtain one-dimensional cooling evaluation results, one-dimensional structural integrity evaluation results, casting feasibility evaluation results, and machining feasibility evaluation results of the turbine blade, and to determine whether the evaluation results are all qualified, and if yes, to enter the determination module, otherwise, to adjust the initial wall thickness of the leading edge, the initial wall thickness of the mid-chord, and the initial wall thickness of the trailing edge of the turbine blade and return to the acquisition module. The determination module is configured to take the initial wall thickness of the leading edge, the initial wall thickness of the mid-chord, and the initial wall thickness of the trailing edge of the turbine blade as the design wall thickness of the leading edge, the design wall thickness of the mid-chord, and the design wall thickness of the trailing edge of the turbine blade.

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