Turbine guide blade leading edge air film cooling structure and application thereof

Abstract

The invention discloses a turbine guide vane front edge air film cooling structure and application thereof, and belongs to the technical field of aero-engine / gas turbine blade cooling. The turbine guide vane front edge air film cooling structure comprises a turbine guide vane front edge, and a through cold air cavity with one closed end is formed in the turbine guide vane front edge; a plurality of conical forming air film cooling holes are formed in the outer side of the front edge of the turbine guide vane in the reverse flow direction or the radial direction along the front edge, the multiple air film cooling holes are sequentially arranged, and one end of each air film cooling hole communicates with the through cold air cavity. The air film cooling effect of the front edge of the turbine guide vane 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 leading edge air film cooling structure of a turbine guide blade and an application thereof. Background Art

[0002] As the core components of aircraft engines and gas turbines, turbine blades undertake the important task of converting the heat energy generated by combustion into mechanical energy. However, this process is accompanied by extremely high temperatures and pressures, which poses a huge challenge to the heat resistance and strength of turbine blade materials. The temperature of the inlet gas before the turbine of an aircraft engine has exceeded 2000K, which is far beyond the heat resistance limit of ordinary metal materials. At such high temperatures, the turbine blade material not only needs to withstand extremely high heat loads, but also faces the impact and corrosion of high-speed airflow. In order to ensure that the high-pressure turbine blades can work safely and reliably in such a harsh environment, a complex cooling structure is designed for the blades. The external air film cooling structure forms a layer of cooling air film on the blade surface to reduce the direct heating of the high-temperature gas on the blade. The internal impingement convection cooling structure introduces cooling air into the blade to reduce the temperature of the blade through convection heat exchange. In addition, a layer of thermal insulation coating is sprayed on the blade surface to further improve the heat resistance and thermal insulation of the blade. This coating usually has low thermal conductivity and high thermal reflectivity, which can effectively reduce the heat transfer from the gas to the blade material. Despite the adoption of a variety of cooling measures, high-pressure turbine blades still face risks such as ablation and cracking. These phenomena not only affect the performance of the blades, but may also cause blade failure, affecting the overall reliability and performance of the engine.

[0003] In order to meet these challenges, researchers and engineers are constantly exploring new cooling technologies and materials. For example, the use of more advanced coating materials, optimization of cooling structure design, and development of new heat-resistant alloys are all hot topics in current research. At the same time, numerical simulation and experimental research are also constantly advancing to better understand and optimize the cooling effect of turbine blades. The development level of high-pressure turbine blade cooling technology is directly related to the technological progress of aircraft engines. Only by solving the cooling problem of turbine blades can the thrust-to-weight ratio, fuel efficiency and reliability of the engine be further improved. Therefore, the research on cooling technology is not only an improvement on the turbine blades themselves, but also a key factor in promoting the development of the entire aircraft engine field. In short, the cooling technology of turbine blades is crucial to the performance and reliability of aircraft engines and gas turbines. With the advancement of science and technology, the cooling technology of turbine blades will be more efficient and reliable in the future, providing strong support for the development of aircraft engines.

[0004] The leading edge of the high-pressure turbine guide vane is the impact stagnation area of ​​the high-temperature combustion gas. The leading edge of the high-pressure turbine guide vane is the first to be directly impacted by the high-temperature and high-pressure combustion gas generated in the combustion chamber. The flow boundary layer has just been formed, and the convective heat transfer intensity is very high. At the same time, the curvature of the leading edge is large, and the cold air has poor wall adhesion, which makes the leading edge area one of the most difficult areas to cool the high-pressure turbine guide vane. In addition, the distribution pattern of the combustion chamber outlet temperature field is high in the middle and low on the upper and lower end walls. The leading edge bears a relatively high heat load. Due to the gas flow characteristics and leading edge geometric characteristics in the guide vane leading edge area, the heat transfer rate of the high-temperature combustion gas toward the wall is relatively high. Therefore, it is very necessary to carry out heat transfer characteristic research on the guide vane leading edge, a key heat transfer area.

[0005] At present, in order to ensure effective cooling of the leading edge, 5 to 7 exhaust film cooling holes are generally set on the leading edge, also known as spray cooling. In the internal chamber of the leading edge, an impingement cooling structure is arranged. After entering the cold air cavity, the cold air first impacts the inner wall of the leading edge to achieve internal cooling, and then flows to the blade surface through the air film cooling holes to form a cooling air film, thereby isolating the high-temperature combustion gas and taking away the heat from the wall. For leading edge spray cooling, how to improve the air film cooling efficiency of the leading edge of the turbine guide vane is a goal that is continuously pursued.

[0006] At present, in order to improve the film cooling efficiency of the leading edge of the turbine guide vane, two technical measures are generally adopted. One is to optimize the leading edge film cooling hole type, such as changing the commonly used cylindrical hole to a fan-shaped hole, a square slot hole or a sister hole. The optimization of the hole structure can significantly improve the film cooling efficiency of the blade. On the other hand, the arrangement of the leading edge film cooling holes is optimized, including the radial angle, flow angle and radial spacing, etc. By optimizing the layout of the film cooling holes, the outflow characteristics of the cold air are improved, and the cooling effect of the cold air is improved. From the current trend of technological development, how to propose a more efficient and practical film cooling hole structure and how to improve the layout of the leading edge film cooling holes are still the key technical approaches that researchers continue to pursue, which has important guiding significance for improving the film cooling effect of the leading edge of the high-pressure turbine guide vanes of aircraft engines. Summary of the invention

[0007] The technical problem to be solved by the present invention is to provide a turbine guide blade leading edge film cooling structure in view of the above-mentioned deficiencies in the prior art, so as to solve the technical problem of poor turbine guide blade leading edge film cooling effect.

[0008] The present invention adopts the following technical solutions: A turbine guide vane leading edge air film cooling structure comprises a turbine guide vane leading edge, a through cold air chamber with one end closed is arranged in the turbine guide vane leading edge, a plurality of air film cooling holes are arranged on the outer side of the turbine guide vane leading edge in the reverse direction or radially along the mainstream flow direction, the plurality of air film cooling holes are arranged in sequence, and one end of each air film cooling hole is connected to the through cold air chamber.

[0009] Preferably, the film cooling hole is a conical structure hole, including a cylindrical segment and an elliptical segment, the diameter D of the cylindrical segment is 0.5-1.2 mm, and the length ratio of the cylindrical segment to the elliptical segment is 2:1.

[0010] Preferably, the rearward tilt angle of the elliptical segment is 7-9°, and the side tilt angle is 7-9°.

[0011] Preferably, the spacing P / D of each exhaust film cooling hole is 4-9.

[0012] Preferably, the film cooling holes include at least 4 to 6 rows of holes.

[0013] Preferably, the film cooling holes in the 3rd and 4th rows are reverse holes, the film cooling holes in the 1st, 2nd, 5th and 6th rows are radial holes, and the film cooling holes in the 2nd, 3rd, 4th and 5th rows are not through.

[0014] Preferably, the film cooling holes in the third and fourth rows are at an angle of 90° to the blade height direction and at an angle of 30-40° to the mainstream flow direction; The film cooling holes in the first, second, fifth and sixth rows have an angle of 20-40° with the blade height direction and an angle of 90° with the mainstream flow direction; The geometric parameters of the first and second rows of film cooling holes are the same as those of the fourth and fifth rows of film cooling holes, and they are arranged opposite to each other along the blade height direction; The film cooling holes in the third and fourth rows are symmetrically distributed along the leading edge stagnation line, and the outlets of the two rows of film cooling holes are located in the range of 20 to 30 degrees on both sides of the stagnation line.

[0015] Preferably, the outlet positions of the film cooling holes in the 1st, 3rd and 5th rows are symmetrical along 50% of the blade height, and the outlet positions of the film cooling holes in the 2nd and 4th rows differ from the outlet positions of the film cooling holes in adjacent rows by 0.5P in the blade height direction.

[0016] Preferably, in the blade height direction, the upper half of the film cooling holes in the 1st, 2nd, 5th and 6th rows are arranged downward in the blade height direction, and the lower half of the film cooling holes in the 1st, 2nd, 5th and 6th rows are arranged upward in the blade height direction.

[0017] 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 film cooling structure.

[0018] Compared with the prior art, the present invention has at least the following beneficial effects: A turbine guide blade leading edge film cooling structure, after cold air is passed through a cold air chamber, it flows out from multiple rows of reverse or radial film cooling holes arranged on the leading edge. Through the improvement of the film cooling hole type and the reasonable organization of the cold air outflow mode, the leading edge can be well covered with cold air to improve the leading edge film cooling effect.

[0019] Furthermore, by designing a conical forming hole, compared with a cylindrical film cooling hole, the outflow momentum of the cold air can be reduced by using the backward expansion and spanwise expansion, while adjusting the outflow path of the cold air, providing better film cooling air wall adhesion performance, and fully improving the cooling effect of the leading edge cold air compared to the cylindrical hole. Compared with the fan-shaped forming hole, the backward-inclined conical forming hole can produce a smaller outlet cross-sectional area and reduce the gas intrusion at the leading edge.

[0020] Furthermore, setting the spacing P / D of each exhaust film cooling hole to 4-9 can effectively ensure that the outflowing cold air of the multiple rows of holes arranged in a staggered row forms good cold air coverage in the radial direction.

[0021] Furthermore, the reverse tapered holes can ensure that within a certain cold air flow range, the cold air passes over the stagnation line to effectively protect the blade surface on the other side, rather than changing the migration path under the action of the mainstream, by symmetrically arranging two rows of reverse outflow film cooling holes within the angle range of 20~30° along the stagnation line, thereby achieving effective protection of the leading edge stagnation zone and improving the film cooling efficiency of the leading edge. For radial film cooling holes, selecting a radial angle of 20~40° can ensure sufficient cold air adherence to the wall and radial expansion.

[0022] Furthermore, the radial counter-arrangement of the film cooling holes can meet the processing requirements in the project and avoid the problem that the designed film cooling solution cannot be processed.

[0023] In summary, the present invention utilizes two rows of oppositely arranged reverse film cooling holes to allow the cold air to cross the leading edge geometric stagnation line to form cold air coverage of the leading edge stagnation zone, thereby reducing the problem of poor wall adhesion of radial film cooling holes caused by the large curvature of the leading edge. The conical film cooling holes are designed to achieve wall adhesion and radial expansion of the reverse hole and radial hole cold air by reducing the outflow momentum of the cold air and enhancing the expansion of the cold air. Compared with the traditional cooling method of arranging cylindrical film cooling holes radially along the leading edge, it can significantly improve the film cooling effect, enhance the leading edge cooling performance under large cold air flow, and improve the film cooling efficiency of the leading edge of the turbine guide vane.

[0024] The technical solution of the present invention is further described in detail below through the accompanying drawings and embodiments. BRIEF DESCRIPTION OF THE DRAWINGS

[0025] Figure 1 This is a schematic diagram of the film cooling structure of the leading edge of the turbine guide blade of the present invention; Figure 2It is a schematic diagram of the arrangement of film cooling holes on the leading edge of the turbine guide vane of the present invention; Figure 3 A schematic diagram of the positional relationship between the cold air cavity and the air film hole at the leading edge of the turbine guide blade of the present invention; Figure 4 It is a front view of the conical film cooling hole of the present invention; Figure 5 It is a side cross-sectional view of the conical film cooling hole of the present invention; Figure 6 It is a side view of the conical film cooling hole of the present invention; Figure 7 A comparison diagram of outflow streamlines of the present invention (left) and the cylindrical radial film cooling hole solution (right); Figure 8 A comparison of the adiabatic film cooling effects of the present invention (left) and the cylindrical radial film cooling hole solution (right); Fig. 9 A comparison of the average film cooling effect at the leading edge of the present invention (left) and the cylindrical radial film cooling hole solution (right).

[0026] Among them: 1. The leading edge of the turbine guide vane; 2. The through cold air chamber; 3. The air film cooling 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 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, 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, unless the context clearly indicates otherwise, the singular forms "a", "an" and "the" are intended to include plural forms.

[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 leading edge air film cooling structure of a turbine guide blade, wherein the multiple air film cooling holes at the leading edge are all supplied with air by a single air supply cavity on the inner side of the leading edge. For the multiple air film cooling holes at the leading edge, the third and fourth air film cooling holes are reverse holes, and the first, second, fifth, and sixth air film cooling holes are all radial holes; the first, second, third, fourth, fifth, and sixth air film cooling holes are all conical structure holes, and the conical structure holes are composed of two sections, a cylindrical section and an elliptical section; by designing the conical structure hole, compared with the cylindrical structure hole, the backward expansion can provide better air film cold air wall adhesion performance; compared with the fan-shaped forming hole, the backward inclined conical forming hole can produce a smaller outlet area, reducing the gas invasion of the leading edge; by designing the reverse conical hole, on the one hand, the wall adhesion of the reverse outflow is improved by the conical hole, and on the other hand, the reverse outflow is used to offset the leading edge stagnation line, ensuring that the leading edge stagnation line area is effectively protected within a certain attack angle range, that is, the air film cooling efficiency is improved, and the cold air coverage at the stagnation line position is ensured at the same time.

[0035] See also Figure 1 The present invention discloses a turbine guide vane leading edge film cooling structure, 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 reversely or radially arranged film cooling holes 3 arranged on the turbine guide vane leading edge.

[0036] The multiple exhaust film cooling 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; 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 interior of the through-type cold air chamber 2, flows through the air film cooling hole 3 channel, and then flows out from the outlet of the multiple exhaust film cooling 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 the guide vane and convection heat exchange, thereby effectively protecting the leading edge 1 of the turbine guide vane and the downstream pressure surface and suction surface.

[0037] Preferably, the film cooling holes 3 are 4 to 6 rows of holes (②③④⑤ or ①②③④⑤⑥).

[0038] If four rows of film cooling holes are provided, two rows of reverse holes are arranged on both sides of the leading edge stagnation line, and then a row of radial holes is arranged downstream of the reverse holes.

[0039] Solution 1: The film cooling holes 3 are 6 rows of holes (①②③④⑤⑥).

[0040] The film cooling holes in the 3rd and 4th rows are reverse holes; The film cooling holes in the 1st, 2nd, 5th and 6th rows are radial holes; The film cooling holes in the 2nd and 3rd rows, and the 4th and 5th rows are not connected to avoid affecting the distribution of cold air flow in the two rows of film cooling holes; The film cooling holes in the third and fourth rows have an angle of 90° with the blade height direction, that is, they are perpendicular to the blade surface in the blade height direction, and the angle with the mainstream flow direction is 30~40°; The film cooling holes in the 1st, 2nd, 5th and 6th rows have an angle of 20-40° with the blade height direction and an angle of 90° with the mainstream flow direction (the tangent direction of the blade surface from the leading edge to the trailing edge), that is, they are perpendicular to the blade surface in the flow direction.

[0041] The geometric parameters of the film cooling holes in the first and second rows are the same as those in the fifth and sixth rows, and they are arranged opposite to each other along the height direction of the blade. The film cooling holes in the third and fourth rows are symmetrically distributed along the leading edge stagnation line, and the outlets of the two rows of film cooling holes are located in the range of 20 to 30 degrees on both sides of the stagnation line.

[0042] Solution 2: The film cooling holes 3 are 4 rows of holes (②③④⑤).

[0043] The film cooling holes in the 3rd and 4th rows are reverse holes; The film cooling holes in the 2nd and 5th rows are radial holes; The film cooling holes in the 2nd and 3rd rows, and the 4th and 5th rows are not connected to avoid affecting the distribution of cold air flow in the two rows of film cooling holes; The film cooling holes in the third and fourth rows have an angle of 90° with the blade height direction, that is, they are perpendicular to the blade surface in the blade height direction, and the angle with the mainstream flow direction is 30~40°; The angle between the 2nd and 5th rows of film cooling holes and the blade height direction is 20~40°, and the angle between them and the mainstream flow direction (the tangent direction of the blade surface from the leading edge to the trailing edge) is 90°, that is, they are perpendicular to the blade surface in the flow direction.

[0044] The geometric parameters of the second row of film cooling holes are the same as those of the fifth row of film cooling holes, and they are arranged opposite to each other along the height direction of the blade; The film cooling holes in the third and fourth rows are symmetrically distributed along the leading edge stagnation line, and the outlets of the two rows of film cooling holes are located in the range of 20 to 30 degrees on both sides of the stagnation line.

[0045] Solution 3: The film cooling holes 3 are 5 rows of holes (①②③④⑤ or ②③④⑤⑥).

[0046] The film cooling holes in the 3rd and 4th rows are reverse holes; The film cooling holes in the 1st, 2nd, 5th and 6th rows are radial holes; The film cooling holes in the 2nd and 3rd rows, and the 4th and 5th rows are not connected to avoid affecting the distribution of cold air flow in the two rows of film cooling holes; The film cooling holes in the third and fourth rows have an angle of 90° with the blade height direction, that is, they are perpendicular to the blade surface in the blade height direction, and the angle with the mainstream flow direction is 30~40°; The film cooling holes in the 1st, 2nd, 5th and 6th rows have an angle of 20-40° with the blade height direction and an angle of 90° with the mainstream flow direction (the tangent direction of the blade surface from the leading edge to the trailing edge), that is, they are perpendicular to the blade surface in the flow direction.

[0047] The geometric parameters of the film cooling holes in the first and second rows are the same as those in the fifth and sixth rows, and they are arranged opposite to each other along the height direction of the blade. The film cooling holes in the third and fourth rows are symmetrically distributed along the leading edge stagnation line, and the outlets of the two rows of film cooling holes are located in the 20-30° interval on both sides of the stagnation line. Film cooling hole 3 is a conical structure hole, including a cylindrical segment and an elliptical segment.

[0048] The diameter D of the cylindrical segment is 0.5-1.2 mm, and the length ratio of the cylindrical segment to the elliptical segment is 2:1; The rear inclination angle of the elliptical segment is 7~9°, and the side inclination angle is 7~9°.

[0049] The spacing P / D of each exhaust film cooling hole 3 is 4-9.

[0050] Among them, the outlet positions of the 1st, 3rd and 5th rows of exhaust film cooling holes are symmetrical along 50% blade height, and the outlet positions of the 2nd and 4th rows of exhaust film cooling holes 3 differ by 0.5P from the outlet positions of the adjacent rows of film cooling holes in the blade height direction.

[0051] Turbine guide blades include turbine guide blades of ground gas turbines and turbine guide blades of aircraft engines.

[0052] The turbine guide vanes of a gas turbine are critical components to ensure efficient engine operation and long life. The design and manufacture of these blades need to take into account the extreme conditions of their operating environment, including high temperatures, high pressures and corrosive gases.

[0053] Nickel-based precipitation hardening equiaxed crystal casting high-temperature alloys such as IN738 or K438 alloy are usually used. These materials have excellent hot corrosion resistance, medium-level high-temperature strength and good organizational stability.

[0054] The blade design needs to take into account aerodynamic performance to ensure smooth airflow into the turbine section. The blade geometry and size need to be precisely controlled to optimize the performance of the engine.

[0055] The manufacturing process usually adopts the investment casting method, which can achieve accurate molding of complex shapes. After casting, heat treatment and surface treatment processes are required to further enhance the heat resistance and wear resistance of the blades.

[0056] Before being put into use, the blades need to undergo rigorous performance testing to ensure their reliability and stability under extreme conditions.

[0057] The turbine guide vanes of aircraft engines need to withstand extreme working conditions, including high temperature, high pressure and high-speed airflow. The design of the blades must take into account both lightness and high strength to ensure efficient operation and long life of the engine. High-temperature alloys such as nickel-based alloys are usually used. These materials have excellent high-temperature strength and resistance to hot corrosion. The blade design needs to consider aerodynamic performance to ensure smooth airflow into the turbine section.

[0058] The geometry and size of the blades need to be precisely controlled to optimize the performance of the engine. They are manufactured using casting or 3D printing technology, which can achieve precise molding of complex shapes. The manufacturing process also includes heat treatment and surface treatment to enhance the heat resistance and wear resistance of the blades. Before being put into use, the blades need to undergo performance testing to ensure their reliability and stability under extreme conditions.

[0059] 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.

[0060] Based on the present invention, an embodiment is given below and compared with a conventional cylindrical radial film cooling solution: In this embodiment, 6 film cooling holes are included, the diameter of the cylindrical section of the film cooling hole is 0.55 mm, the length ratio of the cylindrical section to the elliptical section is 2:1, the rearward inclination angle of the elliptical section is 7°, and the side inclination angle is 7°. The radial angle of the film cooling hole is 25°, the outflow angle of the reverse hole is 30°, and P / D=9.

[0061] As a comparative radial cylindrical hole solution, the cylindrical hole has a diameter of 0.55 mm, a radial angle of 25°, and an outlet position consistent with the embodiment.

[0062] See also Figure 7 , Figure 8 and Fig. 9, by designing reverse holes, the cooling gas can cover the leading edge area of ​​the blade more effectively. This design helps to improve the cooling efficiency, especially in the leading edge stagnation area of ​​the blade; improve the cold air coverage characteristics of the leading edge, ensure that the cooling gas can evenly and effectively cover the entire leading edge area, thereby improving the cooling effect of the blade. Compared with the traditional cylindrical radial holes, this reverse hole design has significantly improved the average air film cooling efficiency along the blade height in the leading edge area, and even exceeds 100% in some local blade height areas. It is suitable for turbine guide blades of ground gas turbines and aircraft engines to improve their performance and life in high temperature, high pressure and corrosive gas environments. Improving cooling efficiency can reduce the thermal stress of blade materials and extend the service life of blades. More efficient cooling can also improve the overall thermal efficiency and performance of the engine. By improving the cooling system of turbine guide blades, the heat resistance and reliability of the blades can be significantly improved, thereby improving the performance and economy of the entire engine.

[0063] In summary, the present invention provides a leading edge film cooling structure of a turbine guide blade. By designing a conical forming hole, compared with a cylindrical film cooling hole, it can provide better film cooling air wall adhesion performance by expanding backward; compared with a fan-shaped forming hole, the backward-inclined conical forming hole can produce a smaller outlet area, reducing the gas intrusion at the leading edge. By designing a reverse conical hole, on the one hand, the wall adhesion of the reverse outflow is improved by the conical hole, and on the other hand, the reverse outflow is used to offset the leading edge stagnation line, ensuring effective protection of the leading edge stagnation line area within a certain range of attack angles, that is, improving the film cooling efficiency, while ensuring cold air coverage at the stagnation line position.

[0064] 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 blade leading edge film cooling structure, characterized in that: The invention comprises a turbine guide vane leading edge (1), a through cold air chamber (2) with one end being closed being arranged in the turbine guide vane leading edge, a plurality of air film cooling holes (3) being arranged on the outer side of the turbine guide vane leading edge in a direction opposite to or radially from the mainstream flow direction, the plurality of air film cooling holes (3) being arranged in sequence, and one end of each air film cooling hole (3) being in communication with the through cold air chamber (2).

2. The turbine guide blade leading edge film cooling structure according to claim 1, characterized in that: The film cooling hole (3) is a conical structure hole, including a cylindrical segment and an elliptical segment, the diameter D of the cylindrical segment is 0.5-1.2 mm, and the length ratio of the cylindrical segment to the elliptical segment is 2:

1.

3. The turbine guide vane leading edge film cooling structure according to claim 2, characterized in that: The rear inclination angle of the elliptical segment is 7~9°, and the side inclination angle is 7~9°.

4. The turbine guide blade leading edge film cooling structure according to claim 1, characterized in that: The spacing P / D of each exhaust film cooling hole (3) is 4-9.

5. The turbine guide vane leading edge film cooling structure according to claim 1, characterized in that: The air film cooling holes (3) include at least 4 to 6 rows of holes.

6. The turbine guide vane leading edge film cooling structure according to claim 5, characterized in that: The air film cooling holes (3) in the third and fourth rows are reverse holes, the air film cooling holes (3) in the first, second, fifth and sixth rows are radial holes, and the air film cooling holes (3) in the second, third, fourth and fifth rows are not through-connected.

7. The turbine guide vane leading edge film cooling structure according to claim 6, characterized in that: The film cooling holes (3) in the third and fourth rows are at an angle of 90° to the blade height direction and at an angle of 30-40° to the mainstream flow direction; The film cooling holes (3) in the first, second, fifth and sixth rows have an angle of 20 to 40 degrees with respect to the blade height direction and an angle of 90 degrees with respect to the mainstream flow direction; The geometric parameters of the first and second rows of exhaust film cooling holes (3) are the same as those of the fourth and fifth rows of exhaust film cooling holes (3), and they are arranged in a hedge relationship along the blade height direction; The film cooling holes in the third and fourth rows are symmetrically distributed along the leading edge stagnation line, and the outlets of the two rows of film cooling holes are located in the range of 20 to 30 degrees on both sides of the stagnation line.

8. The turbine guide vane leading edge film cooling structure according to claim 5, characterized in that: The outlet positions of the film cooling holes (3) in the first, third and fifth rows are symmetrical along 50% of the blade height, and the outlet positions of the film cooling holes (3) in the second and fourth rows differ from the outlet positions of the film cooling holes (3) in the adjacent rows by 0.5P in the blade height direction.

9. The turbine guide vane leading edge film cooling structure according to claim 5, characterized in that: In the blade height direction, the upper parts of the film cooling holes (3) in the first row, the second row, the fifth row and the sixth row are arranged downward in the blade height direction, and the lower parts of the film cooling holes (3) in the first row, the second row, the fifth row and the sixth row are arranged upward in the blade height direction.

10. A high-pressure turbine blade, characterized in that: The leading edge of the blade adopts the turbine guide blade leading edge film cooling structure according to any one of claims 1 to 9.

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