Turbine guider end wall cooling structure and method

Abstract

The invention discloses a turbine guider end wall cooling structure and method.The turbine guider end wall cooling structure comprises a turbine guider end wall, the turbine guider end wall is provided with a turbine guider end wall gas side and a turbine guider end wall cold air side, a ventilation gap is formed between the front end of the turbine guider end wall gas side and a turbine rotor, and the ventilation gap communicates with a cooling gas supply area and a main runner; the cooling gas supply area, the ventilation gap and the main flow channel form a first cooling flow path; the turbine guider support is connected with the turbine guider end wall, a cooling channel is defined by the turbine guider support and the cold air side of the turbine guider end wall, an air inlet hole is formed in the front end of the turbine guider support, and an air outlet hole is formed in the rear end of the turbine guider end wall; the cooling gas supply area, the gas inlet hole, the cooling channel, the gas outlet hole and the main flow channel form a second cooling flow path. A turbine guider end wall front-end air film, front impact, middle strong turbulent flow and tail air film cooling structure is adopted, and the cold air use efficiency can be greatly improved.

Description

Technical Field

[0001] The present invention relates to the field of aero-engines and gas turbines, and particularly to a turbine guide vane end wall cooling structure and method. Background Art

[0002] The turbine guide vane is an important part of the turbine in thermal power devices such as aero-engines and gas turbines. It is usually installed in front of the turbine and mainly consists of structures such as guide vanes. Its function is to guide the high-temperature and high-pressure gas coming out of the combustion chamber so that the gas enters the working blades of the turbine in a certain direction and angle, thereby effectively converting the thermal energy and pressure energy of the gas into the mechanical energy of the turbine, driving the turbine to rotate at a high speed, and then driving components such as the compressor to work.

[0003] The key to the development of advanced aero-engines is to increase the turbine inlet temperature. However, the continuous increase in the turbine inlet temperature requires the development of more advanced cooling technologies. But the focus of people's attention often concentrates on the vane body area of the guide vane, while ignoring the cooling of the guide vane end wall area. On the gas side of the turbine guide vane, generally, the cooling gas of other cooling flow paths flows into the main flow path from the front of the end wall, forming a layer of cold gas film on the end wall surface. On the cold gas side of the turbine guide vane end wall, the cooling gas flowing through other cooling flow paths is used for heat exchange on some smooth walls. The flow direction of the wall surface and the cooling gas is a parallel relationship, or directly flows into the main flow path from some film holes on the end wall. This simple convective cooling structure has poor cooling effect, low cold gas utilization efficiency, and insufficient cold gas utilization efficiency. In the trend of increasing gas temperature, if the traditional turbine guide vane end wall cooling structure is still used, it will cause insufficient end wall cooling and ablation phenomenon. If the cold gas volume is increased, it may cause waste of cold gas use and affect the engine performance. Summary of the Invention

[0004] In view of this, the present invention provides a turbine guide vane end wall cooling structure and method to solve the problem of low cold gas utilization efficiency in the current method for cooling the turbine guide vane end wall.

[0005] In a first aspect, the present invention provides a turbine guide vane end wall cooling structure, including:

[0006] A turbine guide vane end wall, having a turbine guide vane end wall gas side close to the main flow path and a turbine guide vane end wall cold gas side facing away from the main flow path. There is a ventilation gap between the front end of the turbine guide vane end wall gas side and the turbine rotor, and the ventilation gap communicates the cooling gas supply area with the main flow path; the cooling gas supply area, the ventilation gap, and the main flow path form a first cooling flow path;

[0007] The turbine guide vane support is connected to the end wall of the turbine guide vane and forms a cooling channel by surrounding the cold gas side of the end wall of the turbine guide vane. An air inlet hole for connecting the cooling gas supply area and the cooling channel is provided at the front end of the turbine guide vane support, and an air outlet hole for connecting the cooling channel and the main flow channel is opened at the rear end of the end wall of the turbine guide vane; the cooling gas supply area, the air inlet hole, the cooling channel, the air outlet hole and the main flow channel form a second cooling flow path.

[0008] The beneficial effects of the above turbine guide vane end wall cooling method are as follows: The combination of the first cooling flow path and the second cooling flow path constitutes a composite cooling mode of film cooling + impingement cooling + secondary film cooling, breaking through the limitations of traditional single convective cooling, significantly improving the cooling uniformity on the gas side and the cold gas side of the end wall, increasing the utilization rate of cold gas, and effectively suppressing the ablation risk caused by local high temperature.

[0009] In an optional embodiment, a flow disturbing structure is provided in the cooling channel.

[0010] In an optional embodiment, the flow disturbing structure includes a plurality of flow disturbing columns arranged in an array. The flow disturbing columns are provided on the wall surface of the turbine guide vane end wall opposite to the turbine guide vane support, and there is a spacing between the flow disturbing columns and the turbine guide vane support.

[0011] The beneficial effects of the above technical solution are as follows: By adopting the film cooling at the front end of the turbine guide vane end wall + front impingement + strong flow disturbance in the middle + film cooling at the tail structure, a greater improvement in the cold gas usage efficiency can be achieved.

[0012] In an optional embodiment, the cooling channel is sequentially divided into an impingement area and a flow disturbance area from the front end to the rear end. The impingement area and the flow disturbance area are connected through a ventilation hole. Each flow disturbing column is arranged in the flow disturbance area. The air inlet hole is connected to the impingement area, the air outlet hole is connected to the flow disturbance area, and the inner diameter of the ventilation hole is smaller than the inner diameters of the impingement area and the flow disturbance area respectively.

[0013] In an optional embodiment, the aperture H of the air inlet hole is 1 - 1.5 mm, the aperture of the ventilation hole is 0.5H - 0.8H, the circumferential distance between two adjacent flow disturbing columns is 1.5H - 2H, the axial distance between two adjacent flow disturbing columns is 1.5H - 2H, the height of the flow disturbing column is 1.5H - 2H, and the aperture of the air outlet hole is 0.5H - 0.8H.

[0014] The beneficial effects of the above technical solution are as follows: By setting the relative relationship between the cold gas side air inlet aperture of the turbine guide vane end wall and the restricted size of the cooling flow path, the residence time of the cooling gas in the cooling channel is prolonged, and the cooling effect is improved.

[0015] In an alternative embodiment, the intake end of the intake hole is inclined from the rear to the front towards the outlet end of the intake hole, so that after the cooling gas enters the cooling channel, it impacts the end wall of the turbine guide vane, directly cooling the front end of the end wall of the turbine guide vane. And the impact direction forms an obtuse angle with the direction in which the cooling gas flows from the impact area to the turbulent flow area, thereby reducing the flow rate of the cooling gas in the impact area, prolonging the residence time of the cooling gas in the impact area, and improving the heat exchange effect.

[0016] In an alternative embodiment, the intake end of the outlet hole is inclined from the front to the rear towards the outlet end of the outlet hole, so that the cooling gas forms a cooling gas film covering along the rear surface of the end wall of the turbine guide vane, thereby isolating the high-temperature main gas.

[0017] In an alternative embodiment, the turbine rotor extends backward with a protrusion, and the front end of the end wall of the turbine guide vane is located inside the protrusion and forms the ventilation gap with the protrusion. The ventilation gap is inclined from the front to the rear, so that the cooling gas forms a cooling gas film covering along the front surface of the end wall of the turbine guide vane, thereby isolating the high-temperature main gas.

[0018] In a second aspect, the present invention provides a method for cooling the end wall of a turbine guide vane. The method is carried out by using the cooling structure of the end wall of the turbine guide vane, and includes a method for cooling the end wall of the turbine guide vane through a first cooling flow path and a method for cooling the end wall of the turbine guide vane through a second cooling flow path;

[0019] The method for cooling the end wall of the turbine guide vane through the first cooling flow path includes the steps of: the cooling gas flows from the ventilation gap into the main flow channel, and starts to form a cooling gas film covering on the front surface of the gas side of the end wall of the turbine guide vane;

[0020] The method for cooling the end wall of the turbine guide vane through the second cooling flow path includes the steps of: after the cooling gas impacts and cools the front part of the end wall of the turbine guide vane through the intake hole, it is discharged from the outlet hole through the cooling channel, and a cooling gas film is formed on the rear end of the gas side of the end wall of the turbine guide vane.

[0021] In an alternative embodiment, when cooling the end wall of the turbine guide vane through the second cooling flow path, after the cooling gas enters the cooling channel, it is strongly disturbed by the turbulence structure and then discharged from the outlet hole.

[0022] In summary, the technical solution of the present invention has the following advantages:

[0023] The present invention uses a small amount of cold air and has a good cooling effect, can effectively reduce the working temperature of the end wall of the guide vane, and improve the use safety and service life of the turbine guide vane.

[0024] The present invention adopts a cooling structure of film cooling at the front end of the turbine guide vane end wall + front impact + middle strong turbulence + rear film cooling. Among them, impact cooling and turbulator cooling belong to the cooling methods with better heat transfer effects. The heat transfer intensity of impact cooling is about 4 to 6 times that of the smooth wall in the downstream direction, and the heat transfer intensity of turbulator cooling is about 2 to 3 times that of the smooth wall in the downstream direction, which can achieve a large improvement in the cold air utilization efficiency. Brief Description of the Drawings

[0025] In order to more clearly illustrate the specific embodiments of the present invention or the technical solutions in the prior art, the following will briefly introduce the drawings required for use in the description of the specific embodiments or the prior art. Obviously, the drawings in the following description are some embodiments of the present invention. For those of ordinary skill in the art, without creative efforts, other drawings can also be obtained based on these drawings.

[0026] Figure 1 It is a schematic diagram of the cooling flow path of the turbine guide vane end wall cooling structure provided by the present invention;

[0027] Figure 2 It is a schematic diagram of the turbine guide vane end wall cooling structure provided by the present invention;

[0028] Figure 3 It is a schematic diagram of the dimensions of the turbulators of the turbine guide vane end wall cooling structure provided by the present invention.

[0029] Description of the Reference Numerals:

[0030] 1. Turbine rotor, 101. Protrusion, 2. Outer ring of the turbine rotor, 3. Turbine guide vane blade, 4. Turbine guide vane end wall, 41. Gas side of the turbine guide vane end wall, 42. Cold air side of the turbine guide vane end wall, 5. Turbine guide vane bracket, 6. Ventilation gap, 7. Air inlet hole, 8. Ventilation hole, 9. Turbulator, 10. Air outlet hole, 11. Cooling channel, 111. Impact area, 112. Turbulence area, 12. Main flow channel, 13. Cooling gas supply area, 14. Circumferential distance between two adjacent turbulators, 15. Axial distance between two adjacent turbulators. Specific Embodiments

[0031] In order to make the objectives, technical solutions, and advantages of the embodiments of the present invention clearer, the following will clearly and completely describe the technical solutions in the embodiments of the present invention with reference to the drawings in the embodiments of the present invention. Obviously, the described embodiments are some, but not all, of the embodiments of the present invention. Based on the embodiments of the present invention, all other embodiments obtained by those of ordinary skill in the art without creative efforts belong to the scope of protection of the present invention.

[0032] According to an embodiment of the present invention, in a first aspect, a turbine guide vane end wall cooling structure is provided, which is applicable to the turbine guide vane of an aeroengine and a gas turbine rotor, specifically for the end wall of the turbine guide vane, to strengthen the cooling of the end wall of the turbine guide vane and reduce the working temperature of the end wall. In combination with Figures 1 to 3 As shown, the turbine guide vane end wall cooling structure includes a turbine guide vane end wall 4 and a turbine guide vane support 5.

[0033] The turbine guide vane end wall 4 is connected to the vane body 3 of the turbine guide vane. There is a turbine rotor outer ring 2 on the outer side of the turbine rotor 1. The turbine guide vane end wall 4 has a gas side 41 of the turbine guide vane end wall close to the main flow path 12 and a cold gas side 42 of the turbine guide vane end wall facing away from the main flow path 12. There is a ventilation gap 6 between the front end of the gas side 41 of the turbine guide vane end wall and the turbine rotor 1. The ventilation gap 6 communicates the cooling gas supply area 13 with the main flow path 12. The cooling gas supply area 13, the ventilation gap 6 and the main flow path 12 form a first cooling flow path, so that the cooling gas flows into the main flow path 12 from the front end of the turbine guide vane end wall and forms a layer of cooling gas film at the front end of the turbine guide vane end wall.

[0034] The turbine guide vane support 5 is connected to the turbine guide vane end wall 4 and encloses to form a cooling channel 11 with the cold gas side 42 of the turbine guide vane end wall. An air inlet hole 7 is provided at the front end of the turbine guide vane support 5. The air inlet hole 7 is used to communicate the cooling gas supply area 13 with the cooling channel 11. When the cooling gas enters the cooling channel 11 through the air inlet hole 7, it will impact the turbine guide vane end wall 4. An air outlet hole 10 is opened at the rear end of the turbine guide vane end wall 4. The air outlet hole 10 is used to communicate the cooling channel 11 with the main flow path 12. The cooling gas supply area 13, the air inlet hole 7, the cooling channel 11, the air outlet hole 10 and the main flow path 12 form a second cooling flow path, so that after the cooling gas impacts and cools the front part of the turbine guide vane end wall 4 through the air inlet hole 7, it is discharged from the air outlet hole 10 through the cooling channel 11 and forms a layer of cooling gas film at the rear end of the turbine guide vane end wall.

[0035] For the above-mentioned turbine guide vane end wall cooling structure, the combination of the first cooling flow path (the cooling gas forms a front-end gas film through the ventilation gap) and the second cooling flow path (the cooling gas forms a rear-end gas film after impingement cooling) constitutes a composite cooling mode of "film cooling + impingement cooling + secondary film cooling", breaking through the limitations of traditional single convective cooling, significantly improving the cooling uniformity of the gas side and the cold gas side of the end wall, increasing the utilization rate of the cold gas, and effectively suppressing the ablation risk caused by local high temperature. By dividing the cold gas into two cooling flow paths to cool the turbine guide vane end wall 4, the utilization rate of the cold gas is increased.

[0036] The cooling channel 11 is formed by the turbine guide vane support 5 and the turbine guide vane end wall 4, eliminating the need for additional complex cooling cavities, achieving efficient cooling within a limited space, and meeting the lightweight requirements of aero-engines.

[0037] In some embodiments, a turbulator structure is provided in the cooling channel 11. By means of the turbulator structure, the residence time of the cooling gas in the cooling channel 11 is prolonged, the heat transfer intensity is increased, and thus the cooling efficiency is improved.

[0038] The turbulator structure includes a plurality of turbulator posts 9 distributed in an array. The turbulator posts 9 are arranged on the wall surface of the turbine guide vane end wall 4 opposite to the turbine guide vane support 5, and there is a spacing between the turbulator posts 9 and the turbine guide vane support 5.

[0039] In this embodiment, a gas film + front impact + middle strong turbulator + tail gas film cooling structure is adopted at the front end of the turbine guide vane end wall. Among them, impact cooling and turbulator post cooling belong to the cooling methods with better heat transfer effects. The heat transfer intensity of impact cooling is about 4 - 6 times that of the smooth wall in the downstream direction, and the heat transfer intensity of turbulator post cooling is about 2 - 3 times that of the smooth wall in the downstream direction, which can achieve a significant improvement in the utilization efficiency of the cooling air.

[0040] When the cooling gas cools the turbine guide vane end wall 4 through the second cooling flow path, a stepped cooling method is adopted to achieve a stepped utilization of the cooling air energy, avoiding energy waste caused by premature mixing of the cooling air into the main flow channel.

[0041] In some embodiments, the cooling channel 11 is sequentially divided into an impact region 111 and a turbulator region 112 from the front end to the rear end. The impact region 111 and the turbulator region 112 are connected through a vent hole 8. The vent hole 8 is arranged on a partition plate, and the partition plate is arranged on the turbine guide vane end wall 4. Each turbulator post 9 is arranged in the turbulator region 112. The intake hole 7 is connected to the impact region 111, and the outlet hole 10 is connected to the turbulator region 112. The inner diameter of the vent hole 8 is smaller than the inner diameters of both the impact region 111 and the turbulator region 112.

[0042] In some embodiments, the ventilation gap S between the turbine guide vane and the turbine rotor is 1 - 2 mm. The aperture H of the intake hole 7 is 1 - 1.5 mm, the aperture of the vent hole 8 is 0.5H - 0.8H, the circumferential distance 14 between two adjacent turbulator posts is 1.5H - 2H, the axial distance 15 between two adjacent turbulator posts is 1.5H - 2H, the height of the turbulator post 9 is 1.5H - 2H, and the aperture of the outlet hole 10 is 0.5H - 0.8H.

[0043] In this embodiment, by setting the relative relationship between the intake aperture diameter of the turbine guide vane end wall cold air side and the cooling flow path restriction dimension, the residence time of the cooling gas in the cooling channel 11 is extended. The selection of the aperture diameter H of the intake hole 7 enables the cooling gas to enter the impingement area 111 at an appropriate flow rate to form effective impingement cooling. The aperture diameter of the ventilation hole 8 is designed to be 0.5H to 0.8H, which not only ensures that the cooling gas can smoothly flow from the impingement area 111 to the turbulator area 112. The design of the circumferential distance 14 between two adjacent turbulator columns and the axial distance 15 between two adjacent turbulator columns, as well as the height of the turbulator column 9, can maximize the turbulator effect, achieve heat load homogenization, enhance the heat transfer intensity, and at the same time maintain the smoothness of the cooling channel. The design of the aperture diameter of the outlet hole 10 ensures that after the cooling gas fully exchanges heat in the turbulator area 112, it can smoothly flow out of the cooling channel.

[0044] In some embodiments, the intake end of the intake hole 7 is inclined from the rear to the front towards the outlet end of the intake hole 7, so that after the cooling gas enters the cooling channel 11, it impinges on the turbine guide vane end wall 4, directly cooling the front end of the turbine guide vane end wall 4. And the impingement direction forms an obtuse angle with the direction in which the cooling gas flows from the impingement area 111 to the turbulator area 112, thereby reducing the flow rate of the cooling gas in the impingement area 111, extending the residence time of the cooling gas in the impingement area 111, and improving the heat transfer effect.

[0045] In some embodiments, the intake end of the outlet hole 10 is inclined from the front to the rear towards the outlet end of the outlet hole 10, so that the cooling gas forms a cooling gas film along the rear end surface of the turbine guide vane end wall 4, thereby isolating the high-temperature main combustion gas.

[0046] In some embodiments, the turbine rotor 1 extends backward with a protrusion 101, and the front end of the turbine guide vane end wall 4 is located inside the protrusion 101 and forms a ventilation gap 6 with the protrusion 101. The ventilation gap 6 is inclined from the front to the rear, so that the cooling gas forms a cooling gas film along the front end surface of the turbine guide vane end wall 4, thereby isolating the high-temperature main combustion gas.

[0047] In some embodiments, the cold air side 42 of the turbine guide vane end wall extends inward with a first convex ring, and the turbine guide vane support 5 extends outward with a second convex ring. The first convex ring is hermetically clamped into the second convex ring, so that the turbine guide vane support 5 is hermetically clamped and connected to the turbine guide vane end wall 4, and the first convex ring, the cold air side 42 of the turbine guide vane end wall and the side wall of the turbine guide vane support 5 enclose to form the cooling channel 11.

[0048] According to an embodiment of the present invention, in a second aspect, a method for cooling the end wall of a turbine guide vane is provided. This method is carried out by using the end wall cooling structure of the turbine guide vane, and includes a method for cooling the end wall 4 of the turbine guide vane through a first cooling flow path and a method for cooling the end wall 4 of the turbine guide vane through a second cooling flow path.

[0049] Cool the gas side of the end wall of the turbine guide vane through the first cooling flow path: Use the cooling gas flowing through the end wall of the guide vane, that is, the cooling air for cooling the turbine rotor 1. The cooling air flows into the main flow path from the front end of the end wall of the turbine guide vane, forming a layer of cooling gas film covering from the front end of the end wall of the guide vane to isolate the high-temperature main gas. This layer of cooling gas film will continuously mix with the main flow gas as it flows backward, and the cooling effect gradually weakens.

[0050] Cool the cold side of the end wall of the turbine guide vane through the second cooling flow path: Use the cooling gas flowing through the end wall of the guide vane, that is, the cooling air for cooling the turbine rotor 1. After the cooling air impacts and cools the front part of the end wall of the guide vane through the air inlet hole 7 in front of the end wall cover of the guide vane, the cooling gas flows laterally through the ventilation hole 8 of the middle partition of the end wall of the guide vane and enters the spoiler cylinder area. The spoiler cylinder area can strongly disturb the flow of the cooling gas and improve the heat transfer effect. Finally, the cooling gas flows into the main flow path through the air outlet hole 10 at the tail of the end wall of the guide vane, forming a cooling gas film on the gas side at the tail of the end wall of the guide vane to isolate the high-temperature main gas.

[0051] The first cooling flow path forms a cooling gas film on the gas side, effectively isolating the high-temperature main gas and reducing heat radiation and convective heat transfer. The second cooling flow path further strengthens the heat transfer effect between the cold air and the end wall through impingement cooling and strong disturbance in the spoiler cylinder area. The design of the spoiler cylinder area enhances the turbulence of the cooling gas and achieves efficient heat transfer.

[0052] The end wall cooling structure of the turbine guide vane of the present invention has been applied to a certain engine. Through the actual operation proof of the engine and the temperature test experiment, it shows a good cooling effect.

[0053] Although the embodiments of the present invention are described in conjunction with the accompanying drawings, those skilled in the art can make various modifications and variations without departing from the spirit and scope of the present invention. Such modifications and variations all fall within the scope defined by the appended claims.

Claims

1. A turbine guide vane end wall cooling structure, characterized in that: include: A turbine guide vane end wall (4) having a turbine guide vane end wall gas side (41) close to a main flow channel (12) and a turbine guide vane end wall cold air side (42) facing away from the main flow channel (12); a ventilation gap (6) is provided between the front end of the turbine guide vane end wall gas side (41) and the turbine rotor (1); the ventilation gap (6) connects a cooling gas supply region (13) and the main flow channel (12); the cooling gas supply region (13), the ventilation gap (6) and the main flow channel (12) form a first cooling flow path; A turbine guide vane bracket (5) is connected to a turbine guide vane end wall (4) and is arranged with the cooling air side (42) of the turbine guide vane end wall to form a cooling channel (11); the front end of the turbine guide vane bracket (5) is provided with an air inlet hole (7) for connecting a cooling gas supply area (13) with the cooling channel (11); the rear end of the turbine guide vane end wall (4) is provided with an air outlet hole (10) for connecting the cooling channel (11) with a main flow channel (12); the cooling gas supply area (13), the air inlet hole (7), the cooling channel (11), the air outlet hole (10) and the main flow channel (12) form a second cooling flow path.

2. The turbine guide vane end wall cooling structure according to claim 1, characterized in that: A flow disturbance structure is arranged in the cooling channel (11).

3. The turbine guide vane end wall cooling structure according to claim 2, characterized in that: The spoiler structure comprises a plurality of spoiler columns (9) distributed in an array, wherein the spoiler columns (9) are arranged on a wall surface of a turbine guide vane end wall (4) opposite to a turbine guide vane bracket (5), and the spoiler columns (9) are spaced apart from the turbine guide vane bracket (5).

4. The turbine guide vane end wall cooling structure according to claim 3, characterized in that: The cooling channel (11) is divided into an impact area (111) and a spoiler area (112) from the front end to the rear end, the impact area (111) and the spoiler area (112) are connected via a vent hole (8), each spoiler column (9) is arranged in the spoiler area (112), the air inlet hole (7) is connected to the impact area (111), the air outlet hole (10) is connected to the spoiler area (112), and the inner diameter of the vent hole (8) is smaller than the inner diameter of the impact area (111) and the inner diameter of the spoiler area (112).

5. The turbine guide vane end wall cooling structure according to claim 4, characterized in that: The aperture H of the air inlet (7) is 1 to 1.5 mm, the aperture H of the air vent (8) is 0.5H to 0.8H, the circumferential distance (14) between two adjacent spoiler columns is 1.5H to 2H, the axial distance (15) between two adjacent spoiler columns is 1.5H to 2H, the height of the spoiler column (9) is 1.5H to 2H, and the aperture H of the air outlet (10) is 0.5H to 0.8H.

6. The turbine guide vane end wall cooling structure according to any one of claims 1 to 5, characterized in that: The air inlet end of the air inlet hole (7) is arranged to be inclined from the back to the front toward the air outlet end of the air inlet hole (7).

7. The turbine guide vane end wall cooling structure according to any one of claims 1 to 5, characterized in that: The air inlet end of the air outlet hole (10) is arranged to be inclined from front to rear toward the air outlet end of the air outlet hole (10).

8. The turbine guide vane end wall cooling structure according to any one of claims 1 to 5, characterized in that: The turbine rotor (1) is provided with a protrusion (101) extending backwards, the front end of the turbine guide end wall (4) is located inside the protrusion (101) and forms the ventilation gap (6) with the protrusion (101), and the ventilation gap (6) is arranged to be inclined from front to rear.

9. A method for cooling a turbine guide vane end wall, characterized in that: The method is performed using the turbine guide vane end wall cooling structure according to any one of claims 1 to 8, including a method of cooling the turbine guide vane end wall (4) through a first cooling flow path and a method of cooling the turbine guide vane end wall (4) through a second cooling flow path; A method for cooling a turbine guide vane end wall (4) through a first cooling flow path comprises the following steps: cooling gas flows from a ventilation gap (6) into a main flow channel, and a cooling gas film is formed on the front end surface of the turbine guide vane end wall on the gas side (41); A method for cooling a turbine guide vane end wall (4) through a second cooling flow path comprises the following steps: cooling gas is subjected to impact cooling on the front portion of the turbine guide vane end wall (4) through an air inlet hole (7), and then discharged from an air outlet hole (10) through a cooling channel (11), and a cooling gas film is formed at the rear end of the gas side (41) of the turbine guide vane end wall.

10. The method for cooling the turbine guide vane end wall according to claim 9, characterized in that: When the turbine guide vane end wall (4) is cooled through the second cooling flow path, after the cooling gas enters the cooling channel (11), the cooling gas is strongly disturbed by the turbulent structure and then discharged from the outlet hole (10).

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