Turbine guide vane front edge air film cooling structure coping with rotational flow incoming flow influence

Abstract

The turbine guide vane front edge air film cooling structure coping with the rotational flow incoming flow influence comprises a turbine guide vane front edge, a through cold air cavity is formed in the turbine guide vane front edge, and a plurality of rows of air film holes are formed in the outer side of the turbine guide vane front edge at intervals. The air film holes are arranged according to the real isobaric distribution characteristics of the front edge of the turbine guide vane under the influence of rotational flow incoming flow, and one end of each air film hole is communicated with the through cold air cavity. The technical problem that a combustion chamber outlet rotational flow field influences the turbine guide vane front edge gas film cooling characteristic is solved, and the turbine guide vane front edge heat insulation gas film cooling efficiency under the influence of rotational flow incoming flow is improved.

Description

Technical Field

[0001] The invention belongs to the technical field of cooling turbine blades of aircraft engines / gas turbines, and in particular relates to a turbine guide vane leading edge air film cooling structure for coping with the influence of swirl incoming flow. Background Art

[0002] As a critical heat-to-power conversion component in an aircraft engine, the design and cooling technology of turbine blades have a decisive influence on the overall performance and reliability of the engine. The inlet temperature of the high-pressure turbine of an aircraft engine has exceeded 2000 K, which is far beyond the heat resistance limit of ordinary metal materials. In this extreme environment, turbine blades need to have excellent heat resistance and thermal stability. In order to ensure that the high-pressure turbine blades can work safely and reliably at such high temperatures, the blades are designed with complex external film cooling structures and internal impingement convection cooling structures. The design and optimization of these cooling structures are crucial to improving the heat resistance of the blades and extending their service life. 20%~30% of the compressed air of the high-pressure compressor is used for cooling the high-pressure turbine. Although this large amount of cooling air ensures the safety of the turbine blades, it also causes a corresponding loss of engine performance, especially affecting the engine's thrust and fuel efficiency. How to reduce the amount of cooling air used in the engine while ensuring cooling performance is the goal that designers pursue unremittingly. This involves in-depth research and innovative design of the cooling system to achieve a more efficient cooling effect. The design and cooling technology of turbine blades have a crucial impact on the performance and reliability of aircraft engines. Through continuous innovation and optimization, designers are committed to reducing the amount of cooling air used in the engine while ensuring cooling performance, in order to achieve higher efficiency and longer life aircraft engines. Research and development in this field will continue to be a key direction for the advancement of aircraft engine technology.

[0003] With the development of turbine cooling technology, researchers have found that incorporating the effects of the uneven temperature field and swirl field at the combustion chamber outlet into the cooling structure design of the turbine is an effective technical approach to further reduce the amount of turbine cooling air used. By carrying out targeted design of the film cooling structure at the leading edge of the guide vane, the film cooling effect at the leading edge of the guide vane can be effectively improved. This includes optimizing the position, size and shape of the cooling holes, and improving the flow path and distribution of the cooling airflow. Faced with the deterioration of the film cooling effect caused by the swirl field, researchers need to develop new cooling structure design methods to adapt to this complex flow environment. This may involve using advanced computational fluid dynamics (CFD) tools for simulation analysis, as well as conducting a large number of experimental studies to verify the effectiveness of the design scheme.

[0004] In order to reduce the generation of NOx in the combustion chamber, strong swirl is gradually used to stabilize the flame. This strong swirl will form a residual swirl field at the inlet of the high-pressure turbine, and the swirl angle can reach ±45°. The existence of the swirl field has a significant impact on the film cooling effect of the leading edge of the high-pressure turbine guide vane. The swirl field will interfere with the normal flow of the cooling airflow, resulting in a decrease in the film cooling effect and an increase in the heat load of the leading edge area of ​​the blade.

[0005] Therefore, by considering the influence of the swirl field at the combustion chamber outlet and optimizing the film cooling structure design of the turbine guide vane leading edge in a targeted manner, it is expected to significantly improve the cooling effect of the guide vane leading edge, thereby reducing the amount of cooling air used and improving the thermal efficiency and performance of the engine. Summary of the invention

[0006] The technical problem to be solved by the present invention is that in view of the deficiencies in the above-mentioned prior art, based on the study of the influence mechanism of the swirl field at the combustion chamber outlet, a turbine guide vane leading edge air film cooling structure that can cope with the influence of swirl incoming flow is provided, which is used to solve the technical problem that the swirl field at the combustion chamber outlet affects the air film cooling characteristics of the turbine guide vane leading edge, and improve the adiabatic air film cooling efficiency of the turbine guide vane leading edge under the influence of swirl incoming flow.

[0007] The present invention adopts the following technical solutions: A turbine guide vane leading edge air film cooling structure for coping with the influence of swirl flow comprises a turbine guide vane leading edge, a through cold air chamber is arranged inside the turbine guide vane leading edge, a plurality of air film holes are arranged at intervals on the outer side of the turbine guide vane leading edge, the air film holes are arranged along the real isobaric distribution characteristics of the turbine guide vane leading edge under the influence of swirl flow, and one end of each air film hole is connected to the through cold air chamber.

[0008] Preferably, the film holes are arranged along the distribution characteristics of the real isobars of the leading edge of the turbine guide vane under the influence of the swirl flow, specifically: By computational fluid dynamics simulation or experimental test, the pressure field distribution at the leading edge of the guide vane is obtained under the condition of considering the swirl field at the combustion chamber outlet; Based on the pressure field distribution data, the position of the film holes at 50% of the leaf height is determined according to a certain arc length / angle, depending on the number of film holes rows (5 to 7 rows) that need to be arranged on the leading edge and the area where the film holes need to be arranged; Based on the position of the film hole at 50% leaf height, the wall isobar passing through this point is extracted, and the distribution law of the film holes in the leaf height direction is arranged along the isobar.

[0009] Preferably, the projection of the axis of the air film hole on the leading edge profile of the turbine guide vane is parallel to the isobars in the radial direction and is arranged in a counter-hedge arrangement.

[0010] Preferably, the air film holes in the upper half of the leading edge of the turbine guide vane are oriented downward, and the air film holes in the lower half of the leading edge of the turbine guide vane are oriented upward.

[0011] Preferably, the angle between the axis of the air film hole and the height direction of the blade is ±20-35°, and the angle between the axis of the air film hole and the leading edge profile of the blade is 90°.

[0012] Preferably, the air film hole is a conical air film hole, including a cylindrical segment and an elliptical segment, and the ratio of the length of the cylindrical segment to the length of the elliptical segment is 3:2.

[0013] Preferably, the diameter D of the cylindrical segment is 0.45-1.0 mm, and the rearward and side tilt angles of the elliptical segment are 7-9°.

[0014] Preferably, the radial spacing between the film holes at the leading edge of the turbine guide vane is 4D-8D, where D is the diameter of the cylindrical section of the film hole.

[0015] Preferably, 5 to 7 rows of air film holes are arranged at intervals on the leading edge of the turbine guide vane.

[0016] Another technical solution of the present invention is a high-pressure turbine blade, wherein the leading edge of the blade adopts a turbine guide vane leading edge air film cooling structure to cope with the influence of swirl flow.

[0017] Compared with the prior art, the present invention has at least the following beneficial effects: A turbine guide vane leading edge air film cooling structure for coping with the influence of swirl incoming flow. After the high-pressure compressor bleed air enters the through-flow cold air chamber, it flows out from the leading edge surface of the blade through the air film holes, thereby effectively protecting the leading edge, blocking the direct impact of the high-temperature incoming flow on the turbine guide vane, and reducing the leading edge temperature through convective heat exchange, thereby effectively protecting the leading edge of the turbine guide vane and the downstream pressure surface and suction surface. By designing the leading edge air film holes to be distributed along the leading edge isobars under the real swirl incoming flow, the transport effect of the incoming flow angle change caused by the swirl on the leading edge cold air can be improved, while ensuring the reasonable distribution of the leading edge cold air flow on the pressure surface and the suction surface, thereby improving the air film cooling effect of the leading edge under the swirl incoming flow.

[0018] Furthermore, the projection of the axis of the air film hole on the leading edge profile of the turbine guide vane is parallel to the isobars under the influence of the swirl flow in the radial direction, which can ensure the reasonable distribution of the leading edge cold air on the pressure side and the suction side under the swirl flow, and avoid the problem of insufficient cooling on the other side due to the convergence of cold air on the pressure side or the suction side caused by the attack angle of the swirl flow; the air film holes are arranged in a radially opposite manner, which is convenient for processing by laser or electric spark.

[0019] Furthermore, the angle between the axis of the air film hole and the blade height direction is ±20~35°, which can ensure the wall-adherence of the tapered hole air outlet and sufficient expansion in the blade span direction; the angle between the axis of the air film hole and the blade leading edge profile in the flow direction is 90°, which is convenient for processing and ensures the reasonable distribution of the cold air in the leading edge cold air cavity to each exhaust film hole.

[0020] Furthermore, by designing a conical forming hole, compared with a cylindrical air film hole, it is possible to provide better air film cooling air wall adhesion performance by expanding forward.

[0021] Furthermore, the rearward and side tilt angles of the elliptical segment are 7~9 degrees respectively, which can ensure the appropriate momentum of the cold air outflow, avoiding the cold air momentum being too large to directly enter the mainstream, and avoiding excessive expansion to cause structural problems at the leading edge and backflow of high-temperature combustion gas.

[0022] Furthermore, the radial spacing of the film holes on the leading edge of the turbine guide vane is 4D~8D. Too large a spacing is not conducive to the coverage of cold air, and too small a spacing is difficult to process and will affect the selection of the diameter of the film holes.

[0023] Furthermore, arranging 5 to 7 rows of air film holes on the leading edge can ensure effective cooling of the leading edge under high-temperature combustion gas and prevent ablation caused by insufficient cooling.

[0024] To summarize, based on the problem of deterioration of the air film cooling effect of conventional cylindrical air film holes on the leading edge due to the swirl flow, the present invention proposes a guide vane leading edge air film cooling hole type and cooling hole layout scheme that copes with the influence of complex swirl flow. Specifically: by designing the leading edge air film holes to be distributed along the leading edge isobars under the real swirl flow, the transport effect of the leading edge cold air caused by the change in the incoming flow angle caused by the swirl can be improved, while ensuring the reasonable distribution of the leading edge cold air flow on the pressure surface and the suction surface, thereby improving the air film cooling effect of the leading edge under the swirl flow; by designing the cylindrical air film holes as conical forming air film holes, compared with the cylindrical holes, the cold air outflow momentum is reduced and the cold air outflow area is increased, thereby ensuring the wall adhesion of the cold air under swirl transport and enhancing the air film cooling effect of the leading edge under the swirl flow.

[0025] The technical solution of the present invention is further described in detail below through the accompanying drawings and embodiments. BRIEF DESCRIPTION OF THE DRAWINGS

[0026] Figure 1 It is a schematic diagram of the structure of the present invention; Figure 2 It is a three-dimensional diagram of the structure of the present invention; Figure 3 It is a front view of the air film hole of the present invention; Figure 4 It is a cross-sectional view of the air film hole of the present invention; Figure 5 A three-dimensional diagram of the air film hole of the present invention; Figure 6 A comparison of the air film cooling effect of the present invention (middle) and the conventional radial cylindrical air film hole (left) under swirl flow; Figure 7 A comparison diagram of radial distribution of average air film cooling effect at the leading edge of the guide vane corresponding to the present invention and the conventional radial cylindrical air film hole under swirl flow; Among them: 1. The leading edge of the turbine guide vane; 2. The through cold air chamber; 3. The air film hole. DETAILED DESCRIPTION

[0027] The following will be combined with the drawings in the embodiments of the present invention to clearly and completely describe the technical solutions in the embodiments of the present invention. Obviously, the described embodiments are part of the embodiments of the present invention, not all of the embodiments. Based on the embodiments of the present invention, all other embodiments obtained by ordinary technicians in this field without creative work are within the scope of protection of the present invention.

[0028] In the description of the present invention, it should be understood that the terms "center", "longitudinal", "lateral", "up", "down", "front", "back", "left", "right", "vertical", "horizontal", "top", "bottom", "inside", "outside", "one side", "one end", "one side" and the like indicate positions or positional relationships based on the positions or positional relationships shown in the accompanying drawings, and are only for the convenience of describing the present invention and simplifying the description, rather than indicating or implying that the device or element referred to must have a specific orientation, be constructed and operated in a specific orientation, and therefore cannot be understood as limiting the present invention. In addition, the terms "first" and "second" are only used for descriptive purposes, and cannot be understood as indicating or implying relative importance or implicitly indicating the number of technical features indicated. Thus, the features defined as "first" and "second" may explicitly or implicitly include one or more of the features. In the description of the present invention, unless otherwise specified, "multiple" means two or more.

[0029] In the description of the present invention, it should be noted that, unless otherwise clearly specified and limited, the terms "installed", "connected", and "connected" should be understood in a broad sense, for example, it can be a fixed connection, a detachable connection, or an integral connection; it can be a mechanical connection or an electrical connection; it can be a direct connection, or it can be indirectly connected through an intermediate medium, or it can be the internal communication of two components. For ordinary technicians in this field, the specific meanings of the above terms in the present invention can be understood according to specific circumstances.

[0030] It should be understood that when used in this specification and the appended claims, the terms "include" and "comprises" indicate the presence of described features, integers, steps, operations, elements and / or components, but do not exclude the presence or addition of one or more other features, integers, steps, operations, elements, components and / or combinations thereof.

[0031] It should also be understood that the terms used in the present specification are only for the purpose of describing specific embodiments and are not intended to limit the present invention. As used in the present specification and the appended claims, the singular forms "a", "an" and "the" are intended to include plural forms unless the context clearly indicates otherwise.

[0032] It should be further understood that the term "and / or" used in the present description and the appended claims refers to any and all possible combinations of one or more of the associated listed items, and includes these combinations.

[0033] Various structural schematic diagrams of the embodiments disclosed in the present invention are shown in the accompanying drawings. These figures are not drawn to scale, and some details are magnified and some details may be omitted for the purpose of clear expression. The shapes of various regions and layers shown in the figures and the relative sizes and positional relationships therebetween are only exemplary, and may deviate in practice due to manufacturing tolerances or technical limitations, and those skilled in the art may additionally design regions / layers with different shapes, sizes, and relative positions according to actual needs.

[0034] The present invention provides a turbine guide vane leading edge air film cooling structure that copes with the influence of swirl incoming flow. Specifically: by designing the leading edge air film holes to be distributed along the leading edge isobars under the real swirl incoming flow, the transport effect of the leading edge cold air caused by the change in the incoming flow angle caused by the swirl can be improved, while ensuring the reasonable distribution of the leading edge cold air flow on the pressure surface and the suction surface, thereby improving the air film cooling effect of the leading edge under the swirl incoming flow; by designing the cylindrical air film holes as conical forming air film holes, compared with the cylindrical holes, by reducing the cold air outflow momentum and increasing the cold air outflow area, the cold air adhesion to the wall under swirl transport is ensured, and the air film cooling effect of the leading edge under the swirl incoming flow is enhanced.

[0035] The complex swirl flow specifically refers to the residual swirl field formed at the inlet of the high-pressure turbine by the large-scale vortex generated by the swirler in the combustion chamber.

[0036] Turbine guide vanes refer to the first-stage guide vanes of the high-pressure turbine of an aircraft engine.

[0037] See also Figure 1 and Figure 2 The present invention discloses a turbine guide vane leading edge air film cooling structure for coping with the influence of swirl flow, comprising a turbine guide vane leading edge 1, a through cold air chamber 2 arranged in the turbine guide vane leading edge, and a plurality of rows of air film holes 3 arranged on the turbine guide vane leading edge 1, wherein one end of each air film hole 3 is connected to the through cold air chamber 2.

[0038] There are 5 to 7 rows of film holes 3 arranged at intervals on the leading edge 1 of the turbine guide vane. These 5 to 7 rows of film holes 3 are arranged along the distribution characteristics of the real isobars of the leading edge of the turbine guide vane under the influence of the swirl flow. The isobars are obtained by computational fluid dynamics simulation or experimental tests based on pressure sensitive paint. Specifically, by performing computational fluid dynamics simulation on the same blade profile without a cooling structure under the swirl flow condition, the wall pressure distribution of the blade under this condition can be obtained in the post-processing results, and then the isobars of the leading edge of the turbine guide vane can be obtained; or in the cascade test with swirl flow conditions, by applying pressure sensitive paint on the surface of the blade without air cooling, the pressure field distribution of the leading edge of the blade is obtained by taking images, and then the isobars of the leading edge of the turbine guide vane can be obtained.

[0039] The multiple exhaust film holes 3 arranged on the leading edge 1 of the turbine guide vane are all supplied with air by a single through-type cold air chamber 2 on the inner side of the leading edge of the turbine guide vane; the high-pressure air drawn out from the high-pressure compressor is introduced from the bottom (or top) of the through-type cold air chamber 2, enters the through-type cold air chamber 2, flows through the channel of the air film holes 3, and then flows out from the outlets of the multiple rows of air film holes 3, forming a cooling air film on the surface of the leading edge 1 of the turbine guide vane, blocking the direct impact of the high-temperature incoming flow on the turbine guide vane, and reducing the leading edge temperature through convective heat exchange, thereby effectively protecting the leading edge 1 of the turbine guide vane and the downstream pressure surface and suction surface.

[0040] The layout method of the air film hole 3 is: 1) Obtain the leading edge pressure field of the turbine guide vane without air cooling under the influence of complex swirl flow; 2) According to the leading edge profile structure and heat load, determine the number of rows and flow distance of the air film holes 3 on the leading edge 1 of the turbine guide vane, and draw isobars according to the flow distance; 3) The air film holes 3 are arranged according to the leading edge isobaric line, the projection of the axis of the air film holes 3 on the profile surface of the turbine guide vane leading edge 1 is parallel to the isobaric line in the radial direction, and are arranged in a counter-attack manner, the direction of the air film holes 3 in the upper half of the turbine guide vane leading edge 1 is downward, and the direction of the air film holes 3 in the lower half is upward.

[0041] See also Figure 1 and Figure 3 The angle between the axis of the air film hole 3 and the radial direction (blade height direction) is ±20~35°, and the angle between the axis of the air film hole 3 and the blade leading edge profile is 90°, that is, it is perpendicular to the blade surface in the flow direction.

[0042] See also Figure 3 , Figure 4 and Figure 5 The air film holes 3 are all conical air film holes, which include a cylindrical segment and an elliptical segment. The diameter D of the cylindrical segment is 0.45-1.0 mm, the length ratio of the cylindrical segment to the elliptical segment is 3:2, the rear tilt angle of the elliptical segment is 7-9°, the side tilt angle is 7-9°, and the radial spacing P / D of the air film holes is 4-8.

[0043] In order to make the purpose, technical solutions and advantages of the embodiments of the present invention clearer, the technical solutions in the embodiments of the present invention will be clearly and completely described below in conjunction with the drawings in the embodiments of the present invention. Obviously, the described embodiments are part of the embodiments of the present invention, rather than all of the embodiments. The components of the embodiments of the present invention described and shown in the drawings here can usually be arranged and designed in various different configurations. Therefore, the following detailed description of the embodiments of the present invention provided in the drawings is not intended to limit the scope of the claimed invention, but merely represents selected embodiments of the present invention. Based on the embodiments in the present invention, all other embodiments obtained by ordinary technicians in this field without making creative work are within the scope of protection of the present invention.

[0044] See also Figure 6 and Figure 7 Based on the present invention, under the condition of swirl flow at the outlet of an aircraft engine combustion chamber, the film cooling effects of conventional radial cylindrical film holes and the leading edge cooling scheme designed based on the present invention are compared.

[0045] In the experimental design, two different air film hole designs were selected: conventional radial cylindrical air film holes and a leading edge cooling scheme based on the design of the present invention. Both schemes used 6 air film holes with a radial angle of 25° and a hedge arrangement.

[0046] The diameter of the cylindrical section of the cylindrical hole and the conical hole is the same, both 0.55 mm. The radial spacing of the air film holes is 7D, where D is the diameter of the air film hole.

[0047] In the design of the tapered hole, the side inclination angle and the rear inclination angle are both 7°. The cylindrical holes are arranged radially, while the tapered holes are arranged along the leading edge isobars under the swirling flow. This design takes into account the influence of the swirling flow on the film cooling effect.

[0048] Under the condition of swirl flow, the effect of air film cooling will be significantly affected. The swirl field will cause the distribution of cooling airflow to change, affecting the formation and stability of the air film.

[0049] The leading edge cooling scheme designed based on the present invention aims to improve the film cooling effect under swirl flow conditions and reduce the heat load in the leading edge area of ​​the blade by optimizing the layout and design of the film holes.

[0050] By combining experimental research with numerical simulation, a detailed three-dimensional flow field model is established to simulate the film cooling effect under different cooling schemes.

[0051] Evaluation indicators include film cooling efficiency, blade surface temperature distribution, cooling air flow and distribution, etc. These indicators will be used to compare the performance of the two cooling schemes.

[0052] It can be seen from the numerical simulation results that the present invention can achieve the established purpose of the invention. By designing the conical hole, the wall-adhering and expansion effects of the swirl flow of the air film cooling air are guaranteed. By arranging along the isobars, the reasonable distribution of the cooling air on the pressure surface and the suction surface is effectively guaranteed.

[0053] Therefore, the present invention can improve the film cooling effect and cold air distribution effect of the guide vane leading edge under complex swirl flow. Judging from the average cooling effect results of the leading edge, the present invention can increase the radial average film cooling effect of the leading edge by more than 50% compared with the conventional layout.

[0054] In summary, the present invention provides a turbine guide vane leading edge air film cooling structure that copes with the influence of swirl incoming flow. Based on the problem of deterioration of the air film cooling effect of conventional cylindrical air film holes on the leading edge due to swirl incoming flow, the present invention proposes a guide vane leading edge air film cooling hole type and cooling hole layout scheme that copes with the influence of complex swirl incoming flow. Specifically: by designing the leading edge air film holes to be distributed along the leading edge isobars under the real swirl incoming flow, the transport effect of the incoming flow angle change caused by the swirl on the leading edge cold air can be improved, while ensuring the reasonable distribution of the leading edge cold air flow on the pressure surface and the suction surface, thereby improving the air film cooling effect of the leading edge under the swirl incoming flow; by designing the cylindrical air film holes as conical air film holes, compared with cylindrical holes, the cold air outflow momentum is reduced and the cold air outflow area is increased, thereby ensuring the wall adhesion of the cold air under swirl transport and enhancing the air film cooling effect of the leading edge under the swirl incoming flow.

[0055] The above contents are only for explaining the technical idea of ​​the present invention and cannot be used to limit the protection scope of the present invention. Any changes made on the basis of the technical solution in accordance with the technical idea proposed by the present invention shall fall within the protection scope of the claims of the present invention.

Claims

1. A turbine guide vane leading edge film cooling structure for coping with the influence of swirl flow, characterized in that: The invention comprises a turbine guide vane leading edge (1), wherein a through cold air chamber (2) is arranged inside the turbine guide vane leading edge (1), and a plurality of air film holes (3) are arranged at intervals outside the turbine guide vane leading edge (1), wherein the air film holes (3) are arranged along the distribution characteristics of the real isobars of the turbine guide vane leading edge under the influence of the swirl flow, and one end of each air film hole (3) is connected to the through cold air chamber (2).

2. The turbine guide vane leading edge film cooling structure for coping with the influence of swirl flow according to claim 1 is characterized in that: The air film holes (3) are arranged along the distribution characteristics of the real isobars of the leading edge of the turbine guide vane under the influence of the swirl flow, specifically: By computational fluid dynamics simulation or experimental test, the pressure field distribution at the leading edge of the guide vane is obtained under the condition of considering the swirl field at the combustion chamber outlet; Based on the pressure field distribution data, the position of the film holes at 50% of the leaf height is determined according to a certain arc length / angle, depending on the number of film holes rows (5 to 7 rows) that need to be arranged on the leading edge and the area where the film holes need to be arranged; Based on the position of the film hole at 50% leaf height, the wall isobar passing through this point is extracted, and the distribution law of the film holes in the leaf height direction is arranged along the isobar.

3. The turbine guide vane leading edge film cooling structure for coping with the influence of swirl flow according to claim 1 is characterized in that: The projection of the axis of the air film hole (3) on the profile surface of the turbine guide vane leading edge (1) is parallel to the isobars in the radial direction and is arranged in a counter-hedge arrangement.

4. The turbine guide vane leading edge film cooling structure for coping with the influence of swirl flow according to claim 1, characterized in that: The direction of the air film holes (3) in the upper half of the leading edge (1) of the turbine guide vane faces downward, and the direction of the air film holes (3) in the lower half of the leading edge (1) of the turbine guide vane faces upward.

5. The turbine guide vane leading edge film cooling structure for coping with the influence of swirl flow according to claim 1, characterized in that: The included angle between the axis of the air film hole (3) and the blade height direction is ±20-35°, and the included angle between the axis of the air film hole (3) and the blade leading edge profile in the aerodynamic flow direction is 90°.

6. The turbine guide vane leading edge film cooling structure for coping with the influence of swirl flow according to claim 1, characterized in that: The air film hole (3) is a conical air film hole, comprising a cylindrical segment and an elliptical segment, and the ratio of the length of the cylindrical segment to the length of the elliptical segment is 3:

2.

7. The turbine guide vane leading edge film cooling structure for coping with the influence of swirl flow according to claim 5, characterized in that: The diameter D of the cylindrical segment is 0.45~1.0 mm, and the rearward and side tilt angles of the elliptical segment are 7~9°.

8. The turbine guide vane leading edge film cooling structure for coping with the influence of swirl flow according to claim 1, characterized in that: The radial spacing between the air film holes (3) and the leading edge (1) of the turbine guide vane is 4D to 8D, where D is the diameter of the cylindrical section of the air film hole.

9. The turbine guide vane leading edge film cooling structure for coping with the influence of swirl flow according to claim 1, characterized in that: Five to seven rows of air film holes (3) are arranged at intervals on the leading edge (1) of the turbine guide vane.

10. A high-pressure turbine blade, characterized in that: The leading edge of the blade adopts the turbine guide vane leading edge film cooling structure for coping with the influence of the swirl flow as described in any one of claims 1 to 9.

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