Gas film cooling structure for turbine blade leading edge

Abstract

The application discloses a turbine blade leading edge film cooling structure, which comprises a blade leading edge and a plurality of film holes with an equivalent diameter D arranged on the blade leading edge, wherein the plurality of film holes are slot-type holes, and the plurality of film holes comprise a plurality of circumferential film hole rows; wherein in the plurality of circumferential film hole rows, the offset distance S between the center line of each film hole and the stagnation line of the blade leading edge is in a numerical relationship with the equivalent diameter D of the film hole as follows: S = 0.5D-1.0D. The turbine blade leading edge film cooling structure can solve the technical problem that the cooling effect is poor due to the fact that the conventional cylindrical film hole cannot completely cover the blade leading edge with a film.

Description

TECHNICAL FIELD

[0001] The present application belongs to the field of gas turbine cooling technology, and particularly relates to a turbine blade leading edge film cooling structure. BACKGROUND

[0002] The cooling methods of gas turbine turbine blades mainly include film cooling and internal cooling, wherein the film cooling is to extract cooling air from the high-pressure stage of the compressor to the turbine blade, and the cooling air flows out from the small holes densely arranged on the turbine blade, so as to separate the hot gas from the turbine blade to achieve the cooling purpose.

[0003] Due to the complex flow characteristics of the leading edge region of the turbine blade, the conventional cylindrical film holes arranged in the form of row and spray cannot completely cover the leading edge of the blade, especially the region near the stagnation line, and the film cooling effect is low. SUMMARY

[0004] Therefore, the present application provides a turbine blade leading edge film cooling structure to at least partially solve the above technical problems.

[0005] The turbine blade leading edge film cooling structure comprises a blade leading edge and a plurality of film holes with an equivalent diameter D arranged on the blade leading edge, wherein:

[0006] The plurality of film holes are slot-type holes;

[0007] The plurality of film holes comprise a plurality of circumferential film hole rows;

[0008] In the plurality of circumferential film hole rows, the offset distance S between the center line of each film hole and the stagnation line of the blade leading edge and the equivalent diameter D of the film hole satisfy the following numerical relationship: S = 0.5D-1.0D.

[0009] According to the embodiments of the present application, wherein:

[0010] The numerical range of the gas outlet edge angle ψ of each film hole in the plurality of circumferential film hole rows is 0°-90°, wherein the numerical range of the gas outlet edge angle ψ of each film hole in the circumferential film hole row closest to the stagnation line of the blade leading edge is 0°-30°.

[0011] According to the embodiments of the present application, wherein:

[0012] The numerical relationship between the cross-sectional width W of each film hole in the circumferential film hole row closest to the stagnation line of the blade leading edge and the equivalent diameter D of the film hole is: W = 1.4D-2.0D.

[0013] According to the embodiments of the present application, wherein:

[0014] The plurality of circumferential air film hole rows are divided into M air film hole groups, and the N+1th air film hole group is located outside the Nth air film hole group, and the air outlet edge angle of the air film hole in the N+1th air film hole group is greater than the air outlet edge angle of the air film hole in the Nth air film hole group.

[0015] According to the embodiment of the present application, wherein:

[0016] In the M air film hole groups, each air film hole group includes two circumferential air film hole rows, and the two circumferential air film hole rows are symmetrically distributed on both sides of the stagnation line of the blade leading edge.

[0017] According to the embodiment of the present application, wherein:

[0018] In the air film hole group closest to the stagnation line of the blade leading edge, the air film holes in the two circumferential air film hole rows are alternately arranged in staggered positions along the stagnation line.

[0019] According to the embodiment of the present application, wherein:

[0020] In each circumferential air film hole row, the radial spacing P between two air film holes ranges from 5D to 10D.

[0021] According to the embodiment of the present application, wherein:

[0022] The flow direction angle a of each air film hole in the plurality of circumferential air film hole rows ranges from 30° to 90°.

[0023] According to the embodiment of the present application, wherein:

[0024] The radial angle β of each air film hole in the plurality of circumferential air film hole rows ranges from 0° to 60°.

[0025] According to the embodiment of the present application, wherein:

[0026] The transverse expansion angle γ of each air film hole in the plurality of circumferential air film hole rows ranges from 5° to 14°.

[0027] According to the embodiment of the present application, compared with the design method of the air film hole in the related art, the air film hole of the embodiment of the present disclosure adopts a slot type hole different from the traditional cylindrical hole, and the two rows of air film holes closest to the stagnation line are moved circumferentially to the stagnation line, and finally the offset distance S between the hole axis and the stagnation line is kept in a small distance range of 0.5D-1.0D. Compared with the circular hole, because the air outlet edge (long side) width of the slot type hole is larger, the air outlet edge will partially cover the stagnation line area. By arranging the slot type air film hole with a large outlet width near the stagnation line, and cooperating with a small offset distance S, the stagnation line will be in the cold gas coverage area, achieving complete air film coverage of the blade leading edge, improving the cooling effect, and the multiple rows of air film coverage in the leading edge area is uniform, and the air film cooling effect is significantly enhanced. BRIEF DESCRIPTION OF DRAWINGS

[0028] Figure 1 is a schematic diagram of a turbine blade leading edge film cooling structure according to an embodiment of the present application;

[0029] Figure 2A is a schematic diagram of the distribution of film holes on both sides of the stagnation line when the film holes are cylindrical holes;

[0030] Figure 2B is a schematic diagram of the distribution of film holes on both sides of the stagnation line when the film holes are slot-shaped holes according to an embodiment of the present application;

[0031] Figure 3A is a schematic diagram of the circumferential distribution of film holes along the leading edge of the blade when the film holes are cylindrical holes;

[0032] Figure 3B is a schematic diagram of the circumferential distribution of film holes along the leading edge of the blade when the film holes are slot-shaped holes according to an embodiment of the present application;

[0033] Figure 4 is a schematic diagram of the geometry of the film holes according to an embodiment of the present application;

[0034] Figure 5A is a cross-sectional view of the film holes in the blade along the hole width direction according to an embodiment of the present application;

[0035] Figure 5B is a cross-sectional view of the film holes in the blade along the hole length direction according to an embodiment of the present application;

[0036] Figure 6A is a schematic diagram of the coverage of the flow on the surface of the leading edge of the blade when the film holes are cylindrical holes;

[0037] Figure 6B is a schematic diagram of the coverage of the flow on the surface of the leading edge of the blade when the film holes are slot-shaped holes according to an embodiment of the present application. DETAILED DESCRIPTION

[0038] In order to make the objectives, technical solutions and advantages of the present application clearer, the present application is further described in detail below with reference to specific embodiments and the accompanying drawings.

[0039] Hereinafter, embodiments of the present disclosure will be described with reference to the accompanying drawings. It is to be understood, however, that the description is merely exemplary of the present disclosure, and is not intended to limit the scope of the present disclosure. In the following detailed description of the embodiments, numerous specific details are set forth in order to provide a thorough understanding of the embodiments of the present disclosure. However, it would be apparent to one skilled in the art that the various embodiments of the present disclosure can be practiced without these specific details. In other instances, well-known structures and methods are not described in detail in order to avoid obscuring the concepts of the present disclosure.

[0040] The terminology used herein is for the purpose of describing particular embodiments only and is not intended to be limiting of the present disclosure. As used herein, the term "includes" and tautological expressions thereof, such as "including," "includes," "include," "contains," "containing," and so forth, shall be read expansively and without limitation. The terms "comprising," "comprise" and / or "comprised of," and tautological expressions thereof (e.g., "comprising of") will be understood to enable recitations that they do not exclude additional matter.

[0041] All terms used herein (including technical and scientific terms) have the meanings commonly understood by one of ordinary skill in the art unless otherwise defined. It should be noted that the terms used herein are defined as having meanings that are consistent with the context of the specification in which the terms are used and that the terms should not be interpreted in an idealized or overly formal way.

[0042] In the case of using expressions similar to "at least one of A, B, and C, etc.", it is generally to be understood that the expression is to be interpreted to include any of the natural inclusive meanings of the term "or" (i.e., the term "or" is used in the inclusive sense). Similarly, in the case of using expressions similar to "at least one of A, B, or C, etc.", it is to be understood that the expression is to be interpreted to include any of the natural inclusive meanings of the term "or" (i.e., the term "or" is used in the inclusive sense). In the case of using expressions similar to "at least one of A, B, and C, etc.", it is to be understood that the expression is to be interpreted to include any of the natural inclusive meanings of the term "or" (i.e., the term "or" is used in the inclusive sense). In the case of using expressions similar to "at least one of A, B, or C, etc.", it is to be understood that the expression is to be interpreted to include any of the natural inclusive meanings of the term "or" (i.e., the term "or" is used in the inclusive sense).

[0043] In the design of gas turbine, the most direct way to improve the efficiency of gas turbine is to increase the turbine inlet gas temperature, but with the current temperature resistance ability of alloy material, the turbine inlet temperature improvement degree is strongly dependent on the turbine cooling effect. At present, the cooling method of gas turbine turbine blade mainly includes film cooling and internal cooling, among which, the film cooling is to extract cooling air from the high pressure stage of the compressor to the turbine blade, and to flow out from the small holes densely arranged on the turbine blade, so as to separate the hot gas from the turbine blade to achieve the purpose of cooling. Since the cooling gas directly acts on the outer surface of the blade, the cooling efficiency is very high, and it is a commonly used technical means for cooling the turbine blade of the current gas turbine.

[0044] Due to the stagnation effect, the leading edge region of turbine blade usually suffers from very high heat load, and the enhancement of film cooling performance in this region has become the key to improve the overall cooling performance of turbine blade. At present, the leading edge region of turbine blade still uses a large number of conventional cylindrical holes arranged in a row and spray mode. The enhancement of the cooling performance of this region by densely arranging a large number of hole rows essentially relies on the heat removal by a large number of cooling gas outflows, rather than complete film coverage. The main reason is that the cooling near the stagnation line has always been a problem in this field.

[0045] In summary, due to the complex flow characteristics of the leading edge region of turbine blade, the conventional cylindrical film holes arranged in a row and spray mode cannot completely cover the leading edge of the blade, especially the region near the stagnation line, and the film cooling effect is low. Therefore, it is urgent to seek a solution that can improve the cooling effect of the leading edge region of the blade.

[0046] Common methods to improve film cooling effect include: 1. Increasing the amount of cooling gas or increasing the number of film holes, and improving the film cooling effect by densely arranging or large cooling gas; 2. Developing special-shaped holes, and obtaining better film coverage by changing the structure of the film holes; 3. Optimizing the arrangement, and obtaining the best film coverage by cleverly arranging the film holes.

[0047] Based on the design concept of changing the structure of the film holes to obtain better film coverage, the embodiments of the present application provide a turbine blade leading edge film cooling structure to at least partially solve the above technical problems.

[0048] According to the embodiments of the present application,

[0049] The turbine blade leading edge film cooling structure includes a blade leading edge and a plurality of film holes with an equivalent diameter D arranged on the blade leading edge, wherein the plurality of film holes are slot-shaped holes; the plurality of film holes include a plurality of circumferential film hole rows; in the plurality of circumferential film hole rows, the offset distance S between the center line of each film hole and the stagnation line of the blade leading edge and the equivalent diameter D of the film hole satisfy the following numerical relationship: S = 0.5D-1.0D.

[0050] Figure 1 is a schematic view of the turbine blade leading edge film cooling structure according to the embodiments of the present application.

[0051] As Figure 1 shown, the leading edge part in the turbine blade is provided with a plurality of slot-shaped film holes, wherein the plurality of film holes include a plurality of circumferential film hole rows, and the plurality of circumferential film hole rows are divided into M film hole groups, as Figure 1As shown, the six circumferential air film holes are divided into three air film hole groups: the first air film hole group, the second air film hole group, and the third air film hole group. Each air film hole group includes two rows of air film holes, which are symmetrically distributed on both sides of the stabilization line 2.

[0052] like Figure 1 As shown, the leading edge of the turbine blade has multiple groove-shaped film air holes. The center point 1 of the leading edge is the center point of the blade in the direction of blade height, and the stagnation line 2 is the linear area where the flow stagnates in the leading edge region. In the coordinate axis, θ represents the circumferential direction and R represents the radial direction. The multiple groove-shaped film air holes are arranged radially with the center point 1 of the leading edge as a reference, and are arranged symmetrically from top to bottom.

[0053] like Figure 1 As shown, the first film film pore group is the film film pore group closest to the stagnation line 2. The first film film pore group includes the first circumferential film film pore row 3 and the second circumferential film film pore row 4. The first circumferential film film pore row 3 and the second circumferential film film pore row 4 are symmetrically distributed on both sides of the stagnation line 2 at the leading edge of the blade.

[0054] With lag line 2 as the reference center line, the second film film pore group is located outside the first film film pore group; the second film film pore group includes the first second circumferential film film pore row 5 and the second second circumferential film film pore row 6, which are symmetrically distributed on both sides of lag line 2 at the leading edge of the blade.

[0055] The third film film pore group is located outside the second film film pore group. The third film film pore group includes the first third circumferential film film pore row 7 and the second third circumferential film film pore row 8. The first third circumferential film film pore row 7 and the second third circumferential film film pore row 8 are symmetrically distributed on both sides of the lag line 2 at the leading edge of the blade.

[0056] According to an embodiment of this disclosure, further, in the first film film vent group closest to the stagnation line 2, the offset distance S between each film film vent and the stagnation line at the leading edge of the blade (the distance between the central axis of the film film vent and the stagnation line 2) and the equivalent diameter D of the film film vent are related as follows: S = 0.5D - 1.0D. Here, the equivalent diameter D of the film film vent refers to the diameter of a circular vent with the same cross-sectional area as the film film vent.

[0057] Figure 2A This is a schematic diagram showing the distribution of air film holes on both sides of the stagnation line when cylindrical holes are used in related technologies. Figure 2B This is a schematic diagram showing the distribution of air film holes on both sides of the stagnation line when the air film holes are groove-shaped holes according to an embodiment of the present invention. Figure 3A This is a schematic diagram showing the circumferential distribution of film-forming pores at the leading edge of the blade when cylindrical pores are used in related technologies. Figure 3Bis a schematic view of the circumferential distribution of the gas film holes on the leading edge of the blade in the case of the slot-shaped holes according to the embodiment of the present application.

[0058] As shown in Figure 2A , 3A , in the related art, the gas film holes are arranged in the form of cylindrical holes, including multiple rows of gas film holes distributed on both sides of the stagnation line 2, such as the first row of gas film holes 3', the second row of gas film holes 4', the third row of gas film holes 5', the fourth row of gas film holes 6', the fifth row of gas film holes 7', and the sixth row of gas film holes 8'. The multiple rows of gas film holes are arranged at equal intervals in the circumferential direction, and the leading edge stagnation line is generally located in the middle region of the multiple rows of holes.

[0059] As shown in Figure 2B , 3B , compared with the design method of the gas film holes in the related art, the gas film holes in the embodiment of the present application adopt slot-shaped holes that are different from the traditional cylindrical holes, and the two rows of gas film holes closest to the stagnation line are moved in the circumferential direction of the stagnation line, so that the offset distance S between the hole axis and the stagnation line is kept in a small distance range: 0.5D-1.0D.

[0060] Compared with the circular holes, due to the larger width of the gas outlet edge (long side) of the slot-shaped hole, the gas outlet edge will partially cover the stagnation line region. By arranging the slot-shaped gas film holes with a large outlet width near the stagnation line and cooperating with a small offset distance S, the stagnation line will be in the cold gas coverage region, achieving complete gas film coverage of the leading edge of the blade and improving the cooling effect. In the related art, the outlet width of the cylindrical hole is small (only 1D), and if the hole needs to cover the stagnation line, it needs to be very close, which has a machining problem and needs to accurately grasp the position of the stagnation line. The usual design adopts a uniform distribution of multiple rows, and it is difficult to cover the stagnation line. The outlet width of the slot-shaped hole is large (about 3D, which depends on the aspect ratio and expansion angle of the slot-shaped hole), so the two rows of holes on both sides of the stagnation line do not need to be very close, and can form coverage of the stagnation line in the radial direction.

[0061] Figure 4 is a schematic view of the geometric structure of the gas film hole according to the embodiment of the present application.

[0062] As shown in Figure 4 , in the slot-shaped gas film hole, the longer side is the gas outlet edge of the hole, and further, the value of the gas outlet edge angle ψ of each gas film hole in the multiple circumferential rows of gas film holes is in the range of: 0°-90°.

[0063] According to the embodiment of the present application, since the slot hole outlet width is large and the main flow direction in the leading edge flow changes obviously, the outlet edge angle ψ is introduced to define the angle between the hole outlet edge and the circumferential tangent plane. By defining the outlet edge angle ψ, it is convenient to adjust the outlet edge angle of the slot hole according to the actual needs in the design of the air film hole, so as to make the air film cooling air more suitable to the local flow after the angle is adjusted.

[0064] According to the embodiment of the present application, further, in the two circumferential air film hole rows closest to the stagnation line 2, the outlet edge angle ψ of each air film hole ranges from 0° to 30°. The flow characteristics of the leading edge part are that the closer to the stagnation line, the lower the flow velocity, so the smaller the outlet edge angle is set, the stronger the radial flow of the cooling air is, and the more easily the effective coverage is formed. As the hole row is away from the stagnation line, the flow velocity increases and the radial velocity decreases, at this time, the large outlet edge angle is closer to the main flow velocity direction and has a greater effect. Therefore, the outlet edge angle of the air film hole is set in a small angle range, and the two rows of holes closest to the stagnation line in the leading edge cooperate with the small outlet edge angle at the compound angle, forming full coverage of the stagnation line.

[0065] According to the embodiment of the present application, further, with the stagnation line as the reference line, the N+1th air film hole group is located on the outer side of the Nth air film hole group, and the outlet edge angle of the air film hole in the N+1th air film hole group is greater than the outlet edge angle of the air film hole in the Nth air film hole group.

[0066] As shown in Figure 2B With the increase of the flow distance, the outlet edge angles of the subsequent hole rows of the suction surface and the pressure surface gradually increase, for example, the first air film hole group is the air film hole group closest to the stagnation line 2, the outlet edge angles of the air film holes in the first circumferential air film hole row 3 and the second circumferential air film hole row 4 are set to 0°-30°, the second air film hole group is the air film hole group second closest to the stagnation line 2, the outlet edge angles of the air film holes in the first second circumferential air film hole row 5 and the second second circumferential air film hole row 6 are set to 30°-60°, and the third air film hole group is the air film hole group farthest from the stagnation line 2, the outlet edge angles of the air film holes in the first third circumferential air film hole row 7 and the second third circumferential air film hole row 8 are set to 60°-90°.

[0067] According to the embodiment of the present application, the gas film coverage is uniform and the gas film cooling effect of the whole leading edge region is significantly enhanced by gradually increasing the gas film edge angle with the increase of the distance from the stagnation line. According to the flow characteristics of the leading edge region, the closer to the stagnation line, the lower the flow velocity. As the hole row is away from the stagnation line, the streamwise velocity increases and the radial velocity decreases. At this time, the large gas film edge angle is closer to the direction of the main flow velocity. By gradually increasing the gas film edge angle as the hole row is away from the stagnation line, the change of the gas film cooling effect is consistent with the change of the direction of the main flow velocity. The closer to the direction of the main flow, the better the gas film cooling effect. According to the embodiment of the present application, each gas film hole group includes two circumferential gas film hole rows, and the two circumferential gas film hole rows are symmetrically distributed on both sides of the stagnation line of the blade leading edge.

[0068] Further, in the gas film hole group closest to the stagnation line of the blade leading edge, the gas film holes in the two circumferential gas film hole rows are alternately arranged in staggered positions along the stagnation line. By using this arrangement, the intersection of the two circumferential gas film hole rows near the leading edge on the inner wall of the blade can be avoided, and a smaller gas film edge angle is used, which can make the stagnation line be in the cold gas coverage region.

[0069] According to the embodiment of the present application, in each circumferential gas film hole row of the plurality of circumferential gas film hole rows, the radial spacing P between the two gas film holes is set in a reasonable numerical range of 5D-10D.

[0070] Figure 5A is a sectional view of the gas film hole in the blade along the hole width direction according to the embodiment of the present application;

[0071] Figure 5B is a sectional view of the gas film hole in the blade along the hole length direction according to the embodiment of the present application.

[0072] As shown in Figure 5A , 5B , the basic geometric parameters of the slot-shaped gas film hole are defined in the figure, including: the cross-sectional width W, the cross-sectional height H, the hole length L, the straight section length Lt of the hole, the expansion angle γ, the streamwise angle α, the radial angle β, etc. In the circumferential layout design of Figure 3B , the axial position of the multiple rows of gas film holes of the leading edge is determined, and combined with the streamwise angle α and the radial angle β, the slot-shaped hole can be preliminarily positioned.

[0073] Further, the numerical relationship between the cross-sectional width W of each first gas film hole and the equivalent diameter D of the first gas film hole is: W=1.4D-2.0D.

[0074] In order to make the stagnation line all in the cold gas coverage area, the cross-sectional width W (hole long side) of the gas film holes in the two circumferential gas film hole rows in the first gas film hole group needs to be set in a larger range, for example, between 1.4D-2.0D. By arranging the slot-shaped gas film holes with a larger outlet width near the stagnation line and cooperating with a smaller offset distance S, the stagnation line can be made to be in the cold gas coverage area, the blade leading edge is completely covered by the gas film, and the cooling effect is improved.

[0075] According to the embodiment of the present application, the flow direction angle α of each gas film hole in the plurality of circumferential gas film hole rows ranges from 30° to 90°.

[0076] According to the embodiment of the present application, the radial angle β of each gas film hole in the plurality of circumferential gas film hole rows ranges from 0° to 60°, and the radial angle of the gas film hole gradually decreases as the distance from the stagnation line increases. Through this design, the gas flow direction change of the gas film hole is closer to the main flow direction change, so that a better gas film cooling effect is obtained.

[0077] For example, in the first gas film hole group, the radial angle of the gas film holes in the first circumferential gas film hole row 3 and the second circumferential gas film hole row 4 is β=45°, in the second gas film hole group, the radial angle of the gas film holes in the first circumferential gas film hole row 5 and the second circumferential gas film hole row 6 is β=22.5°, and in the third gas film hole group, the radial angle of the gas film holes in the first circumferential gas film hole row 7 and the second circumferential gas film hole row 8 is β=0°.

[0078] According to the embodiment of the present application, the transverse expansion angle γ of each gas film hole in the plurality of circumferential gas film hole rows ranges from 5° to 14°. Considering that the transverse expansion angle γ of the slot-shaped gas film hole further expands the hole outlet width, by setting the transverse expansion angle γ in the above reasonable value range, the gas outlet side of the slot-shaped hole can partially cover the stagnation line area.

[0079] In order to verify the technical effect that the turbine blade leading edge gas film cooling structure adopting the embodiment of the present application can achieve, numerical simulation is performed on the blade leading edge structure with the gas film hole adopting a cylindrical hole and the slot-shaped hole of the embodiment of the present application, Figure 6A and Figure 6B The results of the numerical simulation are shown in the following table. Figure 6A is a schematic diagram of the coverage of the gas flow on the blade leading edge surface when the gas film hole adopts a cylindrical hole; Figure 6B is a schematic diagram of the coverage of the gas flow on the blade leading edge surface when the gas film hole adopts a slot-shaped hole according to the embodiment of the present application.

[0080] In the simulation experiment, the circular hole has a diameter D=1.0 mm, the slot hole has a cross-sectional width W=1.7D and a cross-sectional height H=0.5D. The main flow temperature is 300K, the cold gas temperature is 200K, the main flow velocity is 36m / s, and the cold gas amount of the 6-row air film hole at the leading edge is 1.5% of the main flow.

[0081] According to Figure 6A , It can be seen from the results of Figure 6B that the gas film cooling effect of the turbine blade leading edge gas film cooling structure of the embodiment of the present application is greatly improved compared with the gas film cooling structure in the related art, especially the stagnation line region is effectively covered with airflow, and the uniformity of the multiple rows of air films is significantly improved due to the large outlet width of the slot hole and the reasonable air outlet edge angle. According to the simulation calculation, Figure 6A , the area average gas film cooling effect of the display area in Figure 6B is 0.13 and 0.55 respectively, that is, the cooling effect of the turbine blade leading edge gas film cooling structure of the embodiment of the present application can be improved by more than 2 times compared with the structure in the related art.

[0082] The above-described specific embodiments further illustrate the purpose, technical solutions and beneficial effects of the present application, and it should be understood that the above-described is only a specific embodiment of the present application and is not used to limit the present application, and any modification, equivalent replacement, improvement, etc. made within the spirit and principles of the present application should be included in the protection scope of the present application.

Claims

1. A turbine blade leading edge film cooling structure, comprising a blade leading edge and a plurality of film cooling holes with an equivalent diameter of D disposed on the blade leading edge, wherein: The multiple air film pores are groove-shaped pores; The plurality of air film pores include a plurality of circumferential air film pore rows; In the plurality of circumferential film film pore rows, in the circumferential film film pore row closest to the stabilization line of the leading edge of the blade, the offset distance S between the center line of each film film pore and the stabilization line of the leading edge of the blade, and the equivalent diameter D of the film film pore, are as follows: S = 0.5D - 1.0D. The plurality of circumferential air film holes are arranged into M air film hole groups. With the stagnation line as the reference line, the N+1th air film hole group is located outside the Nth air film hole group, and the air outlet corner of the air film hole in the N+1th air film hole group is larger than the air outlet corner of the air film hole in the Nth air film hole group.

2. The cooling structure according to claim 1, wherein: The range of the outlet angle ψ of each air film hole in the plurality of circumferential air film hole rows is 0°-90°. Among them, the range of the outlet angle ψ of each air film hole in the circumferential air film hole row closest to the stagnation line of the leading edge of the blade is 0°-30°.

3. The cooling structure according to claim 1, wherein: In the circumferential air film pore row closest to the stabilization line of the leading edge of the blade, the numerical relationship between the cross-sectional width W of each air film pore and the equivalent diameter D of the air film pore is: W = 1.4D - 2.0D.

4. The cooling structure according to claim 1, wherein: In the M air film pore groups, each air film pore group includes two circumferential air film pore rows, which are symmetrically distributed on both sides of the lag line at the leading edge of the blade.

5. The cooling structure according to claim 4, wherein: In the air film pore group closest to the stagnation line at the leading edge of the blade, the air film pores in the two circumferential air film pore rows are staggered and alternately arranged along the stagnation line.

6. The cooling structure according to claim 4, wherein: In each of the circumferential air film holes, the radial distance P between two air film holes ranges from 5D to 10D.

7. The cooling structure according to claim 1, wherein: The flow angle α of each air film hole in the plurality of circumferential air film holes ranges from 30° to 90°.

8. The cooling structure according to claim 1, wherein: The radial angle β of each air film hole in the plurality of circumferential air film hole rows ranges from 0° to 60°.

9. The cooling structure according to claim 1, wherein: The lateral expansion angle γ of each air film pore in the plurality of circumferential air film pore rows ranges from 5° to 14°.

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