Estimation Method for the Influence of Gas Film Hole Parameters on the Cold Efficiency Margin of Thermal Barrier Coatings for Turbine Blades

Abstract

The present invention discloses a prediction method for the influence of air film hole parameters on the cold efficiency margin of the thermal insulation coating of a turbine blade. By setting a coating peeling area and a coating covering area on the turbine blade model, changing the aperture or angle of the air film holes on the blade model, and then combining the cooling efficiency formula and the cooling effect margin formula, the cooling effect margin of the turbine blade model under different apertures or angles of the air film holes is obtained. Then, the maximum value of the cooling effect margin of the turbine blade model is obtained through 6σ robust optimization. Through optimization by this method, the cold efficiency margin of the suction surface can be increased to more than 17%, and the cold efficiency margin of the pressure surface can be increased to more than 23%.

Description

Technical Field

[0001] The invention belongs to the technical field of engineering thermophysics, and in particular relates to a method for estimating the influence of air film hole parameters on the cooling efficiency margin of a thermal insulation coating of a turbine blade. Background Art

[0002] Today, high thrust-to-weight ratio aircraft engines have received extensive attention and have become an important development direction. With the continuous improvement of thrust-to-weight ratio, the turbine inlet temperature has increased significantly. Among them, the turbine inlet temperature of aircraft engines with a thrust-to-weight ratio of 15-20 has exceeded 2200℃. In order to adapt to the harsh working environment, the blade protection technologies that are currently widely used are: single crystal, film cooling and thermal barrier coating. However, the operating temperature of the nickel-based single crystal high-temperature alloy with the highest temperature bearing capacity currently used in engineering cannot exceed 1150℃, and film cooling can no longer meet the requirements of blade insulation at this working temperature. In this context, thermal barrier coating is currently the most feasible way to significantly increase the service temperature of turbine blades. Studies have shown that 100-500μm thermal barrier coating can reduce the temperature of the blade substrate by 100:300℃. Thermal barrier coating is a thin layer of oxide ceramic coating applied to high-temperature components at the hot end of aircraft engines, which has properties such as high temperature resistance, low thermal conductivity, and corrosion resistance. As one of the most severe components in gas turbine engines, turbine blades work for a long time in an environment of coupled thermal loads such as high temperature, high-speed rotation, erosion, and corrosion, and are subjected to high-temperature and high-pressure gas, cyclic alternating loads, and centrifugal loads. In addition, the blade geometry, the microstructure and composition of each coating layer and its interface are very complex, which makes the thermal barrier coating coated on the blade surface easy to peel off and fail in this environment. The local shedding of the coating will lead to a decrease in its thermal insulation effect and may cause serious mechanical damage to the blade itself. Film cooling is an important cooling method for protecting blades and their thermal insulation coatings. At present, domestic and foreign researchers' research on film cooling mainly focuses on experiments and numerical simulations. Zhu Huiren et al. used a semi-cylindrical model to experimentally study the film cooling heat transfer and film cooling efficiency of multiple rows of cylindrical holes on the leading edge of the blade, and Liu Yafei conducted a modeling study on the cooling effect and flow loss of typical turbine blades. In terms of numerical simulation, Zhang Ling et al. used a sub-grid scale model to calculate the blade cooling effect of compound-angle film holes, and gave a detailed description of the vortex generated by the combined action of the jet and the mainstream. Bai Chuangjun et al. conducted a numerical simulation on the effect of the film hole angle on the leading edge of the blade on the aerodynamic performance of the turbine. The calculation results show that the cold air injection only has a greater impact on the cooling effect near the cooling hole, and the total pressure loss coefficient increases with the increase of the cold air flow rate.

[0003] Although there are many studies on the failure mechanism of thermal barrier coatings at home and abroad, and the research on film cooling is also quite in-depth. However, there are few studies on the influence of local peeling of thermal barrier coatings on the change of the cooling effect of turbine blades, especially the research on the film hole structure of how the cooling air film weakens the influence brought by this change after the coating peels off. Therefore, it is urgent to propose an optimization method with the highest cooling effect margin before and after the coating peeling failure. Summary of the Invention

[0004] The purpose of the present invention is to improve the cooling effect margin of the turbine blade when local peeling of the thermal insulation coating on the turbine blade occurs, and to provide a method for predicting the influence of film hole parameters on the cooling effect margin of the thermal insulation coating of the turbine blade;

[0005] In order to achieve the above purpose, the present invention adopts the following technical solutions: A method for predicting the influence of film hole parameters on the cooling effect margin of the thermal insulation coating of a turbine blade, including:

[0006] Step 1: Establish a turbine blade model, reduce the first five rows of film holes on the leading edge of the turbine blade model to one row of holes, set the geometric parameter boundary conditions of the film holes, and then set the coating peeling area and the coating covering area on the suction surface and the pressure surface of the turbine blade model. The coating covering area is covered with surface thermal resistance;

[0007] Step 2: Change the aperture or angle of the film holes on the turbine blade model to obtain different film hole parameters, and obtain the outer surface wall temperature of the coating peeling area and the outer surface wall temperature of the coating covering area under different film hole parameters by bilinear interpolation;

[0008] Step 3: According to the outer surface wall temperature of the coating peeling area and the outer surface wall temperature of the coating covering area under different film hole parameters in Step 2, combined with the formula of cooling efficiency, obtain the cooling efficiency of the coating peeling area and the cooling efficiency of the coating covering area under different film hole parameters;

[0009] Step 4: According to the cooling efficiency of the coating peeling area and the cooling efficiency of the coating covering area under different film hole parameters, combined with the formula of cooling effect margin, obtain the cooling effect margin of the turbine blade model under different film hole parameters;

[0010] Step 5: In the suction surface and the pressure surface, repeat Steps 2 to 4 respectively to obtain the cooling effect margin on the suction surface of the turbine blade model under different film hole parameters and the cooling effect margin on the pressure surface of the turbine blade model under different film hole parameters;

[0011] Step 6: Process the cooling effect margin on the suction surface of the turbine blade model under different film hole parameters and the cooling effect margin on the pressure surface of the turbine blade model under different film hole parameters through 6σ robust optimization to obtain the highest cooling effect margin on the suction surface and the highest cooling effect margin on the pressure surface.

[0012] Furthermore, in step 1, the heat insulation effect of the coating coverage area is adjusted by setting different surface thermal resistances.

[0013] Furthermore, the specific steps for obtaining the outer surface wall temperature of the coating peeling area and the outer surface wall temperature of the coating coverage area by bilinear interpolation in step 2 are as follows:

[0014] Step 2.1: Take four groups of film hole parameters around different film hole parameters, which are (α 11 , d 11 ), (α 12 , d 12 ), (α 21 , d 21 ), (α 22 , d 22 ), and measure the corresponding surface wall temperatures in the coating peeling area as T 11 , d 11 , T 12 , d 12 , T 21 , d 21 , T 22 , d 22 are T 11 , T 12 , T 21 , T 22 and the corresponding surface wall temperatures in the coating coverage area as T 31 , T 32 , T 41 , T 42 ;

[0015] Step 2.2: Interpolate (α 11 , d 11 ), (α 12 , d 12 ), (α 21 , d 21 ), (α 22 , d 22 ) by bilinear interpolation, and combine with the surface wall temperatures T 11 , T 12 , T 21 , T 22 in step 2.1 to obtain the outer surface wall temperature T w,e = T 11 ·(1 - μ)·(1 - λ) + T 12 ·(1 - λ)·μ + T 21 ·(1 - μ)·λ + T 22 ·μ·λ, where

[0016] Step 2.3: Interpolate (α 11 , d 11 ), (α 12 , d 12 ), (α 21 , d 21 ), (α 22 , d 22 ) by bilinear interpolation, and combine with the surface wall temperatures T 31 , T 32 , T 41 , T 42 in Step 2.1 to obtain the outer surface wall temperature T w ` ,e = T 31 ·(1 - μ)·(1 - λ) + T 32 ·(1 - λ)·μ + T 41 ·(1 - μ)·λ + T 42 ·μ·λ, where

[0017] Furthermore, substituting the outer surface wall temperature of the coating peeling area in Step 3 into the formula for the cooling efficiency, the formula for the cooling efficiency of the coating peeling area is obtained as:

[0018]

[0019] where, T ∞ and T c are the gas inlet temperature and the cold air inlet temperature respectively, T w,e is the outer surface wall temperature of the coating peeling area, and φ is the cooling efficiency of the coating peeling area;

[0020] And substituting the outer surface wall temperature of the coating covered area into the formula for the cooling efficiency, the formula for the cooling efficiency of the coating covered area is obtained as:

[0021]

[0022] where, T` w,e is the outer surface wall temperature of the coating covered area, and φ` is the cooling efficiency of the coating covered area.

[0023] Furthermore, the formula for the cooling effect margin in Step 4 is:

[0024] η x = φ x - φ` x

[0025] η c = φ c - φ` c

[0026]

[0027] where η x is the cold efficiency difference value after the optimization of the film holes, and φ x and φ` x are the cold efficiencies when the coating peels off and when the coating covers after the optimization of the film holes. η c is the cold efficiency difference value of the film holes under the reference working condition, and φ c and φ` c are the cold efficiencies when the coating peels off and when the coating covers for the film holes under the reference working condition.

[0028] Furthermore, the specific steps of step 6 include:

[0029] Step 6.1: Calculate the average value and the standard deviation

[0030] of the suction surface cooling effect margin according to the cooling effect margin of the suction surface of the blade model under different film hole parameters within the geometric parameter boundary conditions of the film holes. 1_mean_Max and the maximum standard deviation S 2_SD_Max ;

[0031] Step 6.3: Set the weight coefficient W1 of the mean value of the objective function and the weight coefficient W2 of the standard deviation, and W1 + W2 = 1;

[0032] Step 6.4: Set the target value

[0033] Step 6.5: Recalculate the average value μ min and the standard deviation σ min ;

[0034] Step 6.6: According to the obtained S 1_mean_Max 、S 2_SD_Max 、W1, W2, μ min and σ min , combined with 6σ robust optimization, construct the objective function as:

[0035]

[0036] And the average value μ of the suction surface cooling effect margin corresponding to all combined sub-intervals calculated in step 6.5 min and the standard deviation σ min are successively substituted into the objective function, and then the combined sub-interval corresponding to the minimum value of the objective function is selected, and the starting value of the interval of this combined sub-interval is taken as the optimal solution. Then, according to the optimal solution, the cooling effect margin corresponding to the optimal solution is obtained from the cooling effect margins of the suction surface of the blade model at different hole diameters or angles of the film holes, that is, the highest cooling effect margin of the suction surface;

[0037] Step 6.7: Calculate the average value of the pressure surface cooling effect margin according to the cooling effect margins of the pressure surface of the blade model at different hole diameters or angles within the boundary conditions and the standard deviation Repeat steps 6.2 to 6.6 to obtain the highest cooling effect margin of the pressure surface.

[0038] Beneficial effects: In the present invention, a coating peeling area and a coating covering area are set on the turbine blade model, and the hole diameter or angle of the film holes on the blade model is changed. Then, combined with the cooling efficiency formula and the cooling effect margin formula, the cooling effect margins of the turbine blade model at different hole diameters or angles of the film holes are obtained. Then, the maximum value of the cooling effect margin of the turbine blade model is obtained through 6σ robust optimization, so as to predict the values of the angle and hole diameter of the film holes when local peeling of the thermal insulation coating on the turbine blade occurs, which can maximize the cooling effect margin of the blade model, and further improve the cooling effect margin of the turbine blade when local peeling of the thermal insulation coating on the turbine blade occurs. Description of the Drawings

[0039] Figure 1 It is a distribution diagram of the film hole angle.

[0040] Figure 2 It is a schematic diagram of the coating covering area and peeling area on the surface of the turbine blade model.

[0041] Figure 3 It is a schematic diagram of the surface thermal resistance.

[0042] Figure 4 It is a flowchart of 6σ robust optimization.

[0043] Figure 5 It is a flowchart of traversal calculation. Detailed Embodiments

[0044] The following further explains the present invention with reference to the drawings.

[0045] The present invention provides a method for predicting the influence of film hole parameters on the cooling effect margin of the thermal insulation coating of a turbine blade, including:

[0046] Step 1: Establish a turbine blade model, reduce the five rows of film cooling holes on the leading edge of the turbine blade model to one row of holes, set the geometric parameter boundary conditions of the film cooling holes, and then set the coating peeling area and coating covering area on the suction surface and pressure surface of the blade model. The coating covering area is covered with surface thermal resistance.

[0047] Step 2: Change the aperture or angle of the film cooling holes on the turbine blade model to obtain different film cooling hole parameters, and obtain the outer surface wall temperature of the coating peeling area and the outer surface wall temperature of the coating covering area under different film cooling hole parameters by bilinear interpolation.

[0048] Step 3: According to the outer surface wall temperature of the coating peeling area and the outer surface wall temperature of the coating covering area under different film cooling hole parameters in Step 2, combined with the formula of cooling efficiency, obtain the cooling efficiency of the coating peeling area and the cooling efficiency of the coating covering area under different film cooling hole parameters.

[0049] Step 4: According to the cooling efficiency of the coating peeling area and the cooling efficiency of the coating covering area under different film cooling hole parameters, combined with the formula of cooling effect margin, obtain the cooling effect margin of the turbine blade model under different film cooling hole parameters.

[0050] Step 5: In the suction surface and pressure surface, repeat Steps 2 to 4 respectively to obtain the cooling effect margin on the suction surface of the turbine blade model under different film cooling hole parameters and the cooling effect margin on the pressure surface of the turbine blade model under different film cooling hole parameters.

[0051] Step 6: Process the cooling effect margin on the suction surface of the turbine blade model under different film cooling hole parameters and the cooling effect margin on the pressure surface of the turbine blade model under different film cooling hole parameters through 6σ robust optimization to obtain the maximum cooling effect margin of the suction surface and the maximum cooling effect margin of the pressure surface.

[0052] The above steps are as follows:

[0053] In Step 1, establish a turbine blade model by simulation with the specific parameters of the C3X blade. The specific parameters of the C3X blade include the C3X blade profile characteristic parameters, the position and size of the radial cooling holes, and the simulation working conditions of the C3X blade.

[0054] Then reduce the five rows of film cooling holes on the leading edge of the turbine blade model to one row of holes. The distribution of the film cooling holes is as Figure 1 shown, and the geometric parameter boundary conditions of the film cooling holes are:

[0055] Figure 1 The specific parameters of the C3X blade are as follows:

[0056] The C3X blade profile characteristic parameters are:

[0057]

[0058] The position and size of the radial cooling holes are as follows:

[0059] Hole number 1 2 3 x 0.2568 0.3705 0.3503 y 0.0571 0.1278 0.2083 Radius r (mm) 2.35 3.15 3.15 4 5 6 7 0.4650 0.5626 0.6548 0.7360 0.3084 0.4063 0.51162 0.6143 3.15 3.15 3.15 3.15 8 9 10 0.8123 0.8837 0.9523 0.72598 0.8295 0.9397 1.55 1.55 1

[0060] C3X blade simulation conditions:

[0061] Mainstream gas parameters Film cooling parameters Inlet total pressure (Pa) 207360 211300 Inlet total temperature (K) 690 517.5 Inlet turbulence intensity (%) 6.8 -- Outlet static pressure (Pa) 123935 -- Mass flow rate (kg / s) 0.798 <![CDATA[1.71×10 -3 >

[0062] Then, coating peeling areas and coating covering areas are set on the suction surface and pressure surface of the blade model, as Figure 2 shown.

[0063] Finally, a surface thermal resistance is covered in the coating covering area. The surface thermal resistance is as Figure 3 shown. The surface thermal resistance includes fluid or solid units and thin walls. The fluid or solid units are arranged on the wall surface of the thin wall, and its function is to equivalently simulate the heat insulation effect of the thermal barrier coating by setting different surface thermal resistances. And the above settings do not change the geometric structure and size of the turbine blade model. Therefore, the modeling and mesh generation can be greatly simplified, and the influence of the thermal barrier coating and the local peeling of the thermal barrier coating on the blade cooling efficiency can be obtained more quickly under different coating thicknesses and different coating peeling conditions. Among them, the surface thermal resistance thermal conductivity resistance R λ is:

[0064] where Δx is the thickness of the thermal barrier coating and λ is the wall thermal conductivity.

[0065] In this embodiment, the film holes take the blade incidence angle as the initial angle of the film holes, which is set to 0°. When the film holes deflect counterclockwise, it is recorded as positive, and vice versa. And it is stipulated that a 15° deflection of the film hole angle is a simulation calculation point. The angular distribution of all film holes is: And it is stipulated that the film hole distribution is 0.75mm, 0.99mm, 1.25mm. Then all simulation examples are:

[0066]

[0067] Among them, it is stipulated that example 4 is the reference condition.

[0068] In step 2, first, the aperture or angle of the film holes on the turbine blade model is changed to obtain different film hole parameters.

[0069] Then, four groups of film hole parameters are taken from around different film hole parameters, which are (α 11 , d 11 ), (α 12 , d 12 ), (α 21 , d 21 ), (α22 ,d 22 ), and measure (α 11 ,d 11 ), (α 12 ,d 12 ), (α 21 ,d 21 ), (α 22 ,d 22 ). The corresponding surface wall temperatures in the coating peeling area are T 11 , T 12 , T 21 , T 22 and the corresponding surface wall temperatures in the coating covered area are T 31 , T 32 , T 41 , T 42 .

[0070] Next, perform interpolation on (α 11 ,d 11 ), (α 12 ,d 12 ), (α 21 ,d 21 ), (α 22 ,d 22 ). Combine the surface wall temperatures T 11 , T 12 , T 21 , T 22 in step 2.1 to obtain the outer surface wall temperature T w,e = T 11 ·(1 - μ)·(1 - λ) + T 12 ·(1 - λ)·μ + T 21 ·(1 - μ)·λ + T 22 ·μ·λ, where

[0071] Then, perform interpolation on (α 11 ,d 11 ), (α 12 ,d 12 ), (α 21 ,d 21 ), (α 22 ,d 22 ). Combine the surface wall temperatures T 31 , T 32 , T 41 , T 42 in step 2.1 to obtain the outer surface wall temperature T w ` ,e = T 31·(1 - μ)·(1 - λ) + T 32 ·(1 - λ)·μ + T 41 ·(1 - μ)·λ + T 42 ·μ·λ, where

[0072] Finally, substitute different film hole parameters into the outer surface wall temperature T w,e = T 11 ·(1 - μ)·(1 - λ) + T 12 ·(1 - λ)·μ + T 21 ·(1 - μ)·λ + T 22 ·μ·λ of the coating peeling area and the outer surface wall temperature T w ` ,e = T 31 ·(1 - μ)·(1 - λ) + T 32 ·(1 - λ)·μ + T 41 ·(1 - μ)·λ + T 42 ·μ·λ of the coating covered area, and obtain the outer surface wall temperature of the coating peeling area and the outer surface wall temperature of the coating covered area under different film hole parameters.

[0073] In step 3, substitute the outer surface wall temperature of the coating peeling area obtained in step 2 under different film hole parameters into the formula of the cooling efficiency, and the formula of the cooling efficiency of the coating peeling area is:[[]]

[0074]

[0075] where, T ∞ and T c are the gas inlet temperature and the cold air inlet temperature respectively, T w,e is the outer surface wall temperature of the coating peeling area, and φ is the cooling efficiency of the coating peeling area.

[0076] Then substitute the outer surface wall temperature of the coating covered area obtained in step 2 under different film hole parameters into the formula of the cooling efficiency, and the formula of the cooling efficiency of the coating covered area is:[[]]

[0077]

[0078] where, T` w,e is the outer surface wall temperature of the coating covered area, and φ` is the cooling efficiency of the coating covered area.

[0079] Next, through the formula of the cooling efficiency of the coating covered area and the formula of the cooling efficiency of the coating peeling area, the cooling efficiency of the coating peeling area and the cooling efficiency of the coating covered area under different film hole parameters.

[0080] In step 4, substitute the cooling efficiency of the coating stripping area and the cooling efficiency of the coating covering area under different film hole parameters into the cooling effect margin formula to obtain the cooling effect margin of the turbine blade model under different film hole parameters. The cooling effect margin formula is as follows:

[0081] η x =φ x -φ` x

[0082] η c =φ c -φ` c

[0083]

[0084] Where η x is the cold effect difference value after film hole optimization, φ x and φ` x are the cold effect efficiencies when the coating is stripped and when the coating is covered after film hole optimization, η c is the cold effect difference value of the film hole under the reference condition, φ c and φ` c are the cold effects when the coating of the film hole is stripped and when the coating is covered under the reference condition.

[0085] In step 5, in the suction surface and the pressure surface, repeat steps 2 to 4 respectively to obtain the cooling effect margin on the suction surface of the turbine blade model under different film hole parameters and the cooling effect margin on the pressure surface of the turbine blade model under different film hole parameters. This step processes the suction surface and the pressure surface according to the structural characteristics of the turbine blade model, so as to obtain the highest cooling effect margin of the suction surface and the highest cooling effect margin of the pressure surface respectively, and then make the comprehensive cooling effect margin of the whole turbine blade model the highest.

[0086] In step 6, calculate the average value and the standard deviation

[0087] of the cooling effect margin of the suction surface according to the cooling effect margin of the suction surface of the blade model under different film hole parameters within the boundary conditions of the film hole geometric parameters. Then, further divide the angle α parameter interval of each film hole into several angle α sub-intervals of the film hole, and at the same time divide the aperture d parameter interval of each film hole into several aperture d sub-intervals of the film hole. And through the combination of any angle α sub-interval of the film hole and any aperture d sub-interval of the film hole, obtain multiple combined sub-intervals, and then calculate the average value and the standard deviation of the cooling effect margin of the suction surface corresponding to all the combined sub-intervals, and select the maximum average value S 1_mean_Max and the maximum standard deviation S 2_SD_Max .

[0088] Next, set the weight coefficient W1 of the mean value of the objective function and the weight coefficient W2 of the standard deviation, and W1 + W2 = 1, and set the target value And recalculate the average value μ of the suction surface cooling effect margin corresponding to all combined sub-intervals min and the standard deviation σ min .

[0089] Then, according to the obtained S 1_mean_Max 、S 2_SD_Max 、W1, W2, μ min and σ min , combined with 6σ robust optimization, construct the objective function as:

[0090]

[0091] And substitute the average value μ of the suction surface cooling effect margin corresponding to all calculated combined sub-intervals min and the standard deviation σ min into the objective function in turn, then select the combined sub-interval corresponding to the minimum of the objective function, and take the starting value of the interval of this combined sub-interval as the optimal solution. Then, according to the optimal solution, obtain the cooling effect margin corresponding to the optimal solution from the cooling effect margins of the suction surface of the blade model at different orifice diameters or angles of the film holes, that is, the highest cooling effect margin of the suction surface.

[0092] Finally, calculate the average value of the cooling effect margin of the pressure surface according to the cooling effect margins of the pressure surface of the blade model at different orifice diameters or angles within the boundary conditions and the standard deviation Repeat the above steps to obtain the highest cooling effect margin of the pressure surface.

[0093] In this embodiment, the processing process of the above steps on the computer is as Figure 4 shown. Input the minimum orifice diameter d of the film hole and the angle α of the film hole within the geometric parameter boundary conditions of the film hole on the suction surface, and then calculate the cooling effect margin on the suction surface of the turbine blade model at the orifice diameter d and angle α of this film hole through bilinear interpolation, the formula of cooling efficiency, and the formula of cooling effect margin. Then judge whether the orifice diameter d and angle α of this film hole reach the parameter upper boundary. If not, adjust the orifice diameter d and angle α of the film hole and recalculate the cooling effect margin on the suction surface of the turbine blade model at the new orifice diameter d and angle α of the film hole until the orifice diameter d and angle α of the film hole reach the parameter upper boundary. If so, output all the cooling effect margins on the suction surface of the turbine blade model, and calculate the average value of the cooling effect margin of the suction surface according to these cooling effect margins on the suction surface of the turbine blade model and the standard deviation Subdivide the 21 parameter intervals of α into 1000 sub - intervals each, resulting in a total of 21000 sub - intervals. Subdivide the 3 parameter intervals of d in the same way, with each interval divided into 1000 sub - intervals, giving a total of 3000 sub - intervals. Calculate all the combined sub - intervals of the α and d parameters, that is, 21000 * 3000 combined sub - intervals. Input the aperture d of the smallest film hole in each combined sub - interval i and the angle α i , and calculate the average value of the suction surface cooling effect margin within the above - mentioned geometric parameter boundary conditions and the standard deviation in the way described above. Find the maximum average value S 1_mean_Max and the maximum standard deviation S 2_SD_Max from the average values and standard deviations of all combined sub - intervals. Then, set the weight coefficient W1 for the mean value of the objective function and the weight coefficient W2 for the standard deviation, and W1 + W2 = 1, and set the target value After that, recalculate all the combined sub - intervals to obtain the average value μ min of the suction surface cooling effect margin and the standard deviation σ min . According to the obtained S 1_mean_Max 、S 2_SD_Max 、W1、W2、 μ min and σ min , combined with 6σ robust optimization, construct the objective function as follows:

[0094]

[0095] And substitute the average value μ min and the standard deviation σ min of all the recalculated suction surface cooling effect margins into the objective function in turn. Then select the combined sub - interval corresponding to the minimum value of the objective function, and take the starting value of this combined sub - interval as the optimal solution. Then, according to the optimal solution, obtain the cooling effect margin corresponding to the optimal solution from the cooling effect margins of the suction surface of the blade model at different film hole apertures or angles, that is, the highest cooling effect margin of the suction surface. Finally, repeat the above entire process for the pressure surface to obtain the highest cooling effect margin of the pressure surface, where the upper parameter boundary refers to the upper boundary of the geometric parameter boundary conditions of the film hole

[0096] As Figure 5 shown, calculate the maximum value of the cooling effect margin and the corresponding parameters in a traversal manner, and compare with the cooling effect margin corresponding to the optimal solution obtained above to verify the accuracy of the result

[0097] According to the above-verified steps for actual calculation, when the target value is the smallest, the maximum suction surface cooling margin is 17.5204%, and at this time, the geometric parameters of the film holes are the film hole angle and the hole diameter as d = 1.25 mm, and the maximum pressure surface cooling margin is 23.1883%. The corresponding film hole angle and hole diameter are d = 1.25 mm.

[0098] The present invention realizes predicting what values of the angle and hole diameter of the film holes can maximize the cooling effect margin of the blade model when local peeling of the thermal insulation coating on the turbine blade occurs, and then improving the cooling effect margin of the turbine blade when local peeling of the thermal insulation coating on the turbine blade occurs according to the predicted results.

[0099] The above are only the preferred embodiments of the present invention. It should be noted that for those of ordinary skill in the art, without departing from the principle of the present invention, several improvements and refinements can be made, and these improvements and refinements should also be regarded as the protection scope of the present invention.

Claims

1. A method for estimating the influence of film hole parameters on the cooling margin of turbine blade thermal insulation coating, characterized in that: include: Step 1: Establish a turbine blade model, reduce the five film holes on the leading edge of the turbine blade model to one row of holes, set the geometric parameter boundary conditions of the film holes, and then set the coating stripping area and coating covering area on the suction surface and pressure surface of the turbine blade model. The coating covering area is covered with surface thermal resistance; Step 2: Change the aperture or angle of the film hole on the turbine blade model to obtain different film hole parameters, and calculate the outer surface wall temperature of the coating stripping area and the outer surface wall temperature of the coating covering area under different film hole parameters by bilinear interpolation; Step 3: According to the outer surface wall temperature of the coating stripping area and the outer surface wall temperature of the coating covering area under different air film hole parameters in step 2, combined with the cooling efficiency formula, the cooling efficiency of the coating stripping area and the cooling efficiency of the coating covering area under different air film hole parameters are obtained; Step 4: According to the cooling efficiency of the coating stripping area and the cooling efficiency of the coating covering area under different film hole parameters, combined with the cooling effect margin formula, the cooling effect margin of the turbine blade model under different film hole parameters is obtained; Step 5: Repeat steps 2 to 4 on the suction surface and pressure surface respectively to obtain the cooling effect margin on the suction surface of the turbine blade model under different air film hole parameters and the cooling effect margin on the pressure surface of the turbine blade model under different air film hole parameters; Step 6: The cooling effect margin on the suction surface of the turbine blade model under different film hole parameters and the cooling effect margin on the pressure surface of the turbine blade model under different film hole parameters are processed by 6σ robust optimization to obtain the maximum cooling effect margin on the suction surface and the maximum cooling effect margin on the pressure surface; The specific steps of step 6 include: Step 6.1: Calculate the average value μ of the cooling effect margin of the suction surface of the blade model under different film hole parameters within the boundary conditions of the film hole geometry parameters. IE and standard deviation σ IE ; Step 6.2: Subdivide the angle α parameter interval of each air film hole into several air film hole angle α sub-intervals, and subdivide the aperture d parameter interval of each air film hole into several air film hole aperture d sub-intervals, and obtain multiple combined sub-intervals by matching any air film hole angle α sub-interval with any air film hole aperture d sub-interval, and then calculate the average value and standard deviation of the suction surface cooling effect margin corresponding to all combined sub-intervals, and select the maximum average value S 1_mean_Max With the maximum standard deviation S 2_SD_Max ; Step 6.3: Set the weight coefficient W1 of the mean value of the objective function and the weight coefficient W2 of the standard deviation, and W1+W2=1; Step 6.4: Set target value Step 6.5: Recalculate the average value μ of the suction surface cooling effect margin corresponding to all combined sub-intervals in step 6.2 min and standard deviation σ min ; Step 6.6: Based on the obtained S 1_mean_Max , S 2_SD_Max 、W1、W2、 μ min and σ min , combined with 6σ robust optimization, the objective function is constructed as: And the average value μ of the suction surface cooling effect margin corresponding to all combined sub-intervals calculated in step 6.5 is min and standard deviation σ min Substitute the objective function in turn, select the combined sub-interval corresponding to the minimum objective function, and take the starting value of the combined sub-interval as the optimal solution. Then, according to the optimal solution, obtain the cooling effect margin corresponding to the optimal solution from the cooling effect margin of the suction surface of the blade model under different apertures or angles of the air film holes, that is, the maximum cooling effect margin of the suction surface. Step 6.7: Calculate the average value of the pressure surface cooling margin based on the cooling margin of the pressure surface of the blade model under different apertures or angles within the boundary conditions. and standard deviation Repeat steps 6.2 to 6.6 to obtain the maximum cooling effect margin on the pressure surface.

2. The method for estimating the influence of film hole parameters on the cooling margin of turbine blade thermal insulation coating according to claim 1 is characterized in that: In step 1, the heat insulation effect of the coating coverage area is adjusted by setting different surface thermal resistances.

3. The method for estimating the influence of film hole parameters on the cooling margin of turbine blade thermal insulation coating according to claim 1 is characterized in that: The specific steps of calculating the outer surface wall temperature of the coating stripping area and the outer surface wall temperature of the coating covering area by bilinear interpolation in step 2 include: Step 2.1: Take four groups of air film hole parameters from different air film hole parameters, namely (α 11 ,d 11 ),(α 12 ,d 12 ),(α 21 ,d 21 ),(α 22 ,d 22 ), and measure (α 11 ,d 11 ),(α 12 ,d 12 ),(α 21 ,d 21 ),(α 22 ,d 22 ) The corresponding surface wall temperatures in the coating stripping area are T 11 ,T 12 ,T 21 ,T 22 and the corresponding surface wall temperatures in the coating coverage area are T 31 ,T 32 ,T 41 ,T 42 ; Step 2.2: Use bilinear interpolation to find (α 11 ,d 11 ),(α 12 ,d 12 ),(α 21 ,d 21 ),(α 22 ,d 22 ) is interpolated and combined with the surface wall temperature T in step 2.1 11 ,T 12 ,T 21 ,T 22 , the outer surface wall temperature T of the coating stripping area is obtained w,e =T 11 ·(1-μ)·(1-λ)+T 12 ·(1-λ)·μ+T 21 ·(1-μ)·λ+T 22 ·μ·λ, where Step 2.3: Use bilinear interpolation to find (α 11 ,d 11 ),(α 12 ,d 12 ),(α 21 ,d 21 ),(α 22 ,d 22 ) is interpolated and combined with the surface wall temperature T in step 2.1 31 ,T 32 ,T 41 ,T 42 , the outer surface wall temperature T of the coating coverage area is obtained w ` ,e =T 31 ·(1-μ)·(1-λ)+T 32 ·(1-λ)·μ+T 41 ·(1-μ)·λ+T 42 ·μ·λ, where 4. The method for estimating the influence of film hole parameters on the cooling margin of turbine blade thermal insulation coating according to claim 3 is characterized in that: Substituting the outer surface wall temperature of the coating stripping zone in step 3 into the cooling efficiency formula, the cooling efficiency formula of the coating stripping zone is obtained as follows: Among them, T ∞ and T c are respectively the gas inlet temperature and the cold gas inlet temperature, T w,e is the outer surface wall temperature of the coating stripping area, φ is the cooling efficiency of the coating stripping area; Substituting the outer surface wall temperature of the coating coverage area into the cooling efficiency formula, the cooling efficiency formula of the coating coverage area is: Among them, T w ` ,e is the outer surface wall temperature of the coating coverage area, and φ` is the cooling efficiency of the coating coverage area.

5. The method for estimating the influence of film hole parameters on the cooling margin of turbine blade thermal insulation coating according to claim 4 is characterized in that: The cooling effect margin formula in step 4 is: or x =φ x -f x ` or c =φ c -f c ` where η x is the cooling effect difference after optimization of the air film hole, φ x With φ x ` is the cooling efficiency of the coating stripping and coating covering after the air film hole is optimized, η c is the cooling efficiency difference of the air film hole under the reference working condition, φ c With φ c ` is the cold effect of film hole coating peeling and coating covering under the benchmark working condition.

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