Gas turbine end wall film hole layout structure based on pit hole pattern

Abstract

The invention discloses a gas turbine end wall air film hole layout structure based on a pit hole pattern, which applies pit type air film holes with high cooling performance to end wall cooling, performs zoning treatment on an end wall area according to the heat exchange characteristic of the end wall, and further combines the flow characteristic near the end wall to improve the heat exchange efficiency of the end wall. And the position and the outlet angle of the pit type air film hole are adjusted, so that compared with a traditional cylindrical air film hole, the low pneumatic loss, the more uniform air film cold air covering and the higher air film cooling efficiency are achieved. By the adoption of the technical scheme, the area average air film cooling efficiency of the turbine end wall is improved by 5.14-10.22% compared with a conventional layout under the condition that the blowing ratio is 0.5-2.5.

Description

Technical Field

[0001] The invention belongs to the technical field of gas turbine cooling, and in particular relates to a gas turbine turbine end wall air film hole layout structure based on a pit hole type. Background Art

[0002] To improve the output power and efficiency of gas turbines, the turbine inlet temperature needs to be continuously increased, which in turn increases the thermal load on the hot-end components (turbine blades, endwalls, etc.). Currently, the turbine inlet temperature of gas turbines has significantly exceeded the maximum heat resistance temperature of metal materials, necessitating more efficient cooling technologies to ensure safe and stable operation. Hot-end component cooling technologies can be divided into two categories: internal cooling and external cooling. Internal cooling technology involves installing turbulent structures (fins, column-ribs, etc.) or impact holes inside the blades to enhance heat transfer on the internal flow channel wall, thereby achieving the purpose of cooling. External cooling technology focuses on how to better isolate the hot-end component wall from the high-temperature mainstream gas to reduce heat transfer, and mainly includes film cooling and thermal barrier coating technologies. The turbine endwall is exposed to the high-temperature mainstream gas, has a large contact area with the high-temperature mainstream gas, and has a higher thermal load, posing a severe challenge to efficient endwall cooling design.

[0003] At present, the end wall cooling mainly adopts the composite technology of impingement-film cooling. On the one hand, the heat transfer performance on the inner wall surface is increased by the impact effect of the impact hole inside the end wall. On the other hand, the cold air after impingement cooling flows out through the film hole and covers the end wall surface to isolate the high-temperature combustion gas. In the current end wall cooling structure design, cylindrical film holes with low processing and manufacturing difficulty are usually used, and the film holes are evenly arranged along the span of the end wall. This conventional arrangement has the following problems: (1) Cylindrical film holes are very easy to blow away from the wall surface under high blowing ratio, resulting in poor cooling effect on the turbine end wall surface, making the cooling efficiency low; (2) The uniform arrangement of film holes along the span leads to uneven cooling distribution on the turbine end wall surface, resulting in overcooling of some areas and poor cooling of other areas, especially in the area of ​​the end wall surface close to the pressure surface where the cold air cannot cover. The high cooling efficiency characteristics of the pit-type film hole have been applied in flat film cooling under cavity inlet and cross-flow conditions. At present, this high-efficiency hole type has not been used in turbine end wall cooling. Summary of the Invention

[0004] The technical problem to be solved by the present invention is to provide a gas turbine turbine end wall air film hole layout structure based on a pit hole type, apply high-efficiency pit-type air film holes to the cooling design of the turbine end wall, and specifically design the layout of the hole type on the end wall for the complex flow near the wall surface of the turbine end wall to improve the cooling efficiency of the turbine end wall cooling structure.

[0005] To achieve the above object, the present invention adopts the following technical solutions:

[0006] A gas turbine turbine end wall air film hole layout structure based on a pit hole type is characterized in that zoning is performed according to the heat exchange characteristics of the turbine end wall surface, and multiple pit-type air film holes are arranged in the end wall area close to the pressure surface of the turbine blade. The pit-type air film holes provide a flow channel for cold air to cool the turbine end wall, and cold air used for air film cooling is provided through a cold air supply cavity. The outlet angle of the pit-type air film holes is adjusted according to the flow characteristics near the turbine end wall surface, so as to achieve higher air film cooling efficiency and uniform cold air coverage effect.

[0007] Preferably, the pit-type air film hole includes a cylindrical part and a pit part, the cylindrical part is a cylindrical air film hole, the pit part is obtained when the thermal barrier coating is sprayed on the turbine end wall, and the cross-section of the pit part remains unchanged along the normal direction of the end wall surface.

[0008] Preferably, the turbine end wall is divided into region A, region B, region C, and region D according to the heat transfer characteristics of the turbine end wall; wherein, region A is located in the leading edge area close to the suction side, region B is located between the horseshoe vortex on the pressure side and the horseshoe vortex on the suction side, region C is close to the pressure side, and region D is located downstream of the end wall close to the trailing edge.

[0009] Preferably, one pit-type air film hole is arranged on the suction side and the pressure side of region A respectively; three pit-type air film holes are arranged along the span direction at the low pressure gradient in region B to utilize the high pressure gradient downstream of region B to cool the downstream wall surface; two pit-type air film holes are arranged along the pressure side at the low pressure gradient in region C; six pit-type air film holes are staggered along the span direction in region D to cool the end wall surface with a high heat transfer coefficient; a total of 13 pit-type air film holes are arranged on the end wall surface.

[0010] Preferably, the outlet angle of each concave-shaped air film hole is determined according to the direction of the near-wall streamline at the location of the air film hole, and the outlet direction is tangent to the streamline.

[0011] Preferably, the geometric parameters of the pit portion are optimized and designed in a simplified double-arc curved channel.

[0012] As a preference, a simplified double-arc curved channel is obtained based on the pressure surface and suction surface profiles of the gas turbine turbine end wall, and the pressure gradient and secondary lateral flow characteristics of the simplified double-arc curved channel wall are kept similar to those of the blade channel.

[0013] Preferably, the gas turbine turbine end wall is the first stage stationary blade or moving blade end wall.

[0014] In order to solve the problems of low cooling efficiency and uneven spatial distribution under high blowing ratio in the conventional end wall film cooling hole layout, the layout structure of the present invention introduces pit-type film holes with high cooling performance to replace the cylindrical film holes in the conventional design. The area is partitioned according to the heat transfer characteristics of the end wall surface, and multiple film holes are arranged in the end wall area close to the pressure surface of the turbine blade which is difficult to cool. The outlet angle of the pit-type film holes is adjusted according to the flow characteristics near the turbine end wall surface, so as to achieve higher film cooling efficiency and more uniform cold air coverage. BRIEF DESCRIPTION OF THE DRAWINGS

[0015] In order to more clearly illustrate the embodiments of the present invention or the technical solutions in the prior art, the following briefly introduces the drawings required for use in the embodiments or the description of the prior art. Obviously, the drawings described below are merely embodiments of the present invention. For ordinary technicians in this field, other drawings can be obtained based on the provided drawings without paying any creative work.

[0016] Figure 1 This is the overall diagram of the layout structure of the gas film holes on the turbine end wall;

[0017] Figure 2 is the geometrical size of the new concave air film hole;

[0018] Figure 3 This is the layout diagram of the air film holes on the end wall;

[0019] Figure 4 is the streamline diagram of the end wall;

[0020] Figure 5 This is the lateral pressure gradient cloud diagram when there is no air film hole on the end wall;

[0021] Figure 6 is the heat transfer coefficient cloud diagram of the end wall;

[0022] Figure 7 The cloud diagram of film cooling efficiency when the blowing ratio is 1.0;

[0023] Figure 8 The area-averaged film cooling efficiency curve for different blowing ratios. DETAILED DESCRIPTION

[0024] The following will clearly and completely describe the technical solutions in the embodiments of the present invention in conjunction with the accompanying drawings. Obviously, the described embodiments are only 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 making creative efforts are within the scope of protection of the present invention.

[0025] In order to make the above-mentioned objects, features and advantages of the present invention more obvious and easy to understand, the present invention is further described in detail below with reference to the accompanying drawings and specific embodiments.

[0026] Example 1:

[0027] like Figure 1 As shown, an embodiment of the present invention provides a gas turbine turbine endwall film hole layout structure based on a concave hole pattern. This structure is arranged in the endwall region near the pressure surface of the turbine blade. It is characterized by the fact that, in the mainstream direction (X-axis), the mainstream airflow flows from the main inlet and flows through the mainstream channel to the mainstream outlet. To achieve effective cooling of the endwall, multiple concave film holes are arranged on the endwall surface. Cooling air enters through the cold air inlet in the vertical direction (Z-axis direction) and is supplied to the concave film holes through a cold air supply cavity located below the film holes, providing an effective cooling air film for the turbine endwall. The concave film holes consist of a cylindrical portion and a concave portion. The cylindrical portion is a conventional cylindrical hole, while the concave portion is formed when the thermal barrier coating is sprayed on the endwall surface. Its cross-section remains unchanged along the normal direction of the endwall surface. Based on the local three-dimensional flow characteristics near the turbine endwall surface, the arrangement density and outlet angle of the concave film holes are optimized, thereby ensuring high film cooling efficiency while improving the uniformity of cold air coverage and adhesion.

[0028] As an implementation method of an embodiment of the present invention, the turbine end wall is divided into regions A, B, C, and D according to the heat transfer characteristics of the turbine end wall; wherein, region A is located in the leading edge near the suction side area, and has a larger heat transfer coefficient due to the stagnation of the mainstream flow and the generation and development of the horseshoe vortex on the suction side; region B is located between the horseshoe vortex on the pressure side and the horseshoe vortex on the suction side, and the flow is relatively stable, so that the heat transfer coefficient of this area is relatively low; region C is close to the pressure side, and has a medium heat transfer coefficient due to the lateral flow caused by the pressure gradient and the development of the secondary flow vortex system; region D is located downstream of the end wall near the trailing edge area, and the air flow velocity in this area is relatively high, and has a higher heat transfer coefficient. The distribution of the number of film holes refers to the fact that the number of film holes arranged in each partition on the turbine end wall depends on the heat transfer characteristics of the partition. When the heat transfer coefficient of the partition is large, multiple film holes need to be arranged, and when the heat transfer coefficient of the partition is small, a smaller number of film holes are arranged. Specifically, one film hole is placed on the suction side and one on the pressure side of region A. Three film holes are arranged along the spanwise direction at a low pressure gradient in region B to cool the downstream wall using the high pressure gradient downstream of region B. Two film holes are placed along the pressure side at a low pressure gradient in region C. Six film holes are staggered along the spanwise direction in region D to cool the endwall surface with a high heat transfer coefficient. A total of 13 film holes are arranged on the endwall surface. The exit angle of each pit-shaped film hole is determined by the direction of the near-wall streamline at the location of the film hole, and its exit direction is tangent to the streamline.

[0029] As an implementation method of an embodiment of the present invention, the geometric parameter design of the pit portion is performed by determining the value range of the optimized geometric parameters, combining Latin hypercube sampling to determine the sample parameters, and using the initial samples to perform neural network training to obtain a proxy model. The proxy model is then optimized using a genetic algorithm to obtain the values ​​of the optimized geometric parameters. The end wall of the annular cascade of the GE-E3 first-stage high-pressure turbine stator blade is used as the research object. The pressure surface and suction surface profile lines at the root position of the turbine blade are stretched to obtain a simplified double-arc curved channel. The pressure gradient and secondary lateral flow characteristics of the simplified double-arc curved channel wall are kept similar to those of the cascade channel. The turbine end wall of a gas turbine is the end wall of the first-stage stator or moving blade.

[0030] Unlike conventional cylindrical film holes, which are uniformly distributed on the end wall, this embodiment of the present invention utilizes high-performance concave film holes for end wall cooling. The end wall area is partitioned based on the heat transfer characteristics of the end wall. Furthermore, the position and exit angle of the concave film holes are adjusted based on the flow characteristics near the end wall. This results in lower aerodynamic losses, more uniform film cooling coverage, and higher film cooling efficiency compared to conventional cylindrical film holes. By replacing the film hole pattern and optimizing the layout, this invention improves the area-averaged film cooling efficiency of the turbine end wall by 5.14-10.22% compared to conventional layouts at a blowing ratio of 0.5-2.5.

[0031] like Figure 2 As shown, it is a new type of concave-type air film hole. The figure shows the spatial layout of the air film hole in a three-dimensional coordinate system, where the X-axis represents the mainstream direction, the Y-axis is the lateral direction perpendicular to the mainstream direction, and the Z-axis represents the direction perpendicular to the end wall. In this design, the diameter of the air film hole is D, and the jet angle is 30° (that is, the angle with the Z-axis is 30°). This hole type is based on a concave-type air film hole. By simulating the end wall channel in a curved channel, the air film hole is optimized. The final design shows the best cooling effect as shown in the figure. Compared with ordinary cylindrical holes in curved channels, this hole type has a significant increase in both the air film coverage area and the coverage distance.

[0032] like Figure 3 The figure shows the lateral pressure gradient cloud diagram when there is no air film hole at the end wall. The chord length C of the turbine blade structure is 62.2 mm, and the axial chord length C axIt is 36.2mm, and the grating pitch P is 47.8mm. In the end wall channel, due to the restriction of fluid flow, a high pressure gradient area may be formed. This is caused by the acceleration or constraint of the fluid in the narrow channel. Therefore, the layout of the air film holes should try to ensure the uniform distribution of the air film in the end wall channel to provide balanced gas pressure support. The layout density of the air film holes needs to be reasonably selected according to the specific application scenarios and requirements. Too high a density may increase the difficulty of manufacturing and maintenance, while too low a density may lead to uneven air film support effect. At the same time, the layout of the air film holes should take into account the influence of fluid flow in the end wall channel. Especially in the high pressure gradient area, it is necessary to ensure that the layout of the air film holes will not have an adverse effect on the flow, such as increasing resistance or interfering with flow stability.

[0033] Figure 4 The figure shows the distribution of convective heat transfer coefficients at the end wall, dividing the end wall into four sections. Within the end wall channels, restricted fluid flow may create regions of high heat transfer coefficients. Depending on the heat flux distribution, it may be necessary to increase the density of film holes in these areas of high heat flux to improve convective heat transfer.

[0034] Figure 5 The figure shows the extreme streamline distribution at the end wall. The streamline density at the end wall is relatively high, which may form a dense or sparse streamline organization. In addition, due to the influence of the surface resistance of the object, the streamline will show a large degree of curvature, so the direction of the streamline is difficult to predict.

[0035] Figure 6 The figure shows the layout of the film holes on the end wall of the present invention. The placement of film holes requires comprehensive consideration of various factors. This solution is based on a comprehensive consideration of the flow characteristics of the end wall, the transverse pressure gradient contour, the convective heat transfer coefficient, the distribution of limiting streamlines, and other factors. The film holes are densely packed in high-heat-transfer areas, reducing placement in high-pressure-gradient areas. The film holes are always aligned parallel to the limiting streamlines.

[0036] Figure 7 The figure shows the cooling efficiency cloud diagram of the non-optimized hole type and layout structure and the layout structure of the present invention on the end wall. It can be seen from the figure that the present invention effectively improves the film cooling efficiency of each high heat load area of ​​the end wall, and the film cooling effect downstream of the end wall and the entire end wall is also significantly improved, and the cold air coverage is more uniform. Figure 8 Comparing the data shown, the new film holes effectively enhance film cooling efficiency, performing well at all air blowing ratios, and performing even better as the air blowing ratio increases. The new hole layout not only provides more uniform cooling coverage on the end wall, but also increases the cooling efficiency of the film holes.

[0037] In summary, the layout structure of the present invention can more evenly distribute the cooling gas on the end wall, enhancing the cooling effect. A reasonable layout of the film holes ensures that the cooling gas effectively covers the entire end wall surface and forms a uniform air film, providing better thermal protection, reducing temperature gradient differences, and avoiding the concentration of local thermal stress, thereby extending the service life of the material. Furthermore, the film hole injection angle, span angle, and film hole configuration can be flexibly adjusted according to actual conditions to further optimize the overall film cooling efficiency.

[0038] Compared with the existing turbine air film hole uniform layout technology based on cylindrical air film holes, the present invention has the following advantages:

[0039] 1. The shape of the air film hole used in the present invention is a geometric shape that has been optimized in a curved channel. This hole shape shows a high cooling coverage area and cooling efficiency on the simulated end wall.

[0040] 2. The present invention improves the cooling efficiency and uniformity of cold air coverage in the end wall area by arranging holes based on the flow characteristics and convective heat transfer features of the end wall, while reducing damage to the overall structure of the end wall, enhancing the structural strength of the end wall, and optimizing the aerodynamic characteristics of the blade.

[0041] 3. The present invention can be applied to gas turbines and aircraft engines, and the injection angle and span angle of the film holes can be adjusted according to actual conditions to optimize the film cooling efficiency of the end wall, meet the cooling requirements of different end walls, and make the overall design more universal.

[0042] The embodiments described above are merely descriptions of preferred embodiments of the present invention and are not intended to limit the scope of the present invention. Without departing from the spirit of the present invention, various modifications and improvements made to the technical solutions of the present invention by persons skilled in the art should fall within the scope of protection defined by the claims of the present invention.

Claims

1. A gas turbine end wall film hole layout structure based on a pit hole type, characterized in that: The turbine end wall is zoned according to its heat transfer characteristics, and a plurality of concave-shaped film holes are arranged in the end wall area close to the pressure surface of the turbine blade. The concave-shaped film holes provide a flow channel for cold air to cool the turbine end wall, and cold air for film cooling is provided through the cold air supply cavity. The outlet angle of the concave-shaped film holes is adjusted according to the flow characteristics near the turbine end wall to achieve higher film cooling efficiency and uniform cold air coverage.

2. The gas turbine end wall film hole layout structure based on the pit hole type according to claim 1 is characterized in that: The pit-type air film hole includes a cylindrical part and a pit part. The cylindrical part is a cylindrical air film hole. The pit part is obtained when the thermal barrier coating is sprayed on the turbine end wall. The cross section of the pit part remains unchanged along the normal direction of the end wall surface.

3. The gas turbine end wall film hole layout structure based on the pit hole type according to claim 2, characterized in that: According to the heat transfer characteristics of the turbine end wall, it is divided into region A, region B, region C, and region D; among them, region A is located in the leading edge area close to the suction side, region B is located between the horseshoe vortex on the pressure side and the horseshoe vortex on the suction side, region C is close to the pressure side, and region D is located downstream of the end wall close to the trailing edge.

4. The gas turbine end wall film hole layout structure based on the pit hole type according to claim 3 is characterized in that: One concave-shaped air film hole is arranged on the suction side and one on the pressure side of region A. Three concave-shaped air film holes are arranged along the span direction at the low pressure gradient in region B to utilize the high pressure gradient downstream of region B to cool the downstream wall. Two concave-shaped air film holes are arranged along the pressure surface at the low pressure gradient in area C; In region D, six concave-shaped air film holes are staggered along the span direction to cool the end wall surface with a high heat transfer coefficient; a total of 13 concave-shaped air film holes are arranged on the end wall surface.

5. The gas turbine end wall film hole layout structure based on the pit hole type according to claim 4 is characterized in that: The outlet angle of each concave-shaped air film hole is determined according to the direction of the near-wall streamline at the location of the air film hole, and its outlet direction is tangent to the streamline.

6. The gas turbine end wall film hole layout structure based on the pit hole type according to claim 5, characterized in that: The geometric parameters of the pit part are optimized and designed in a simplified double-arc curved channel.

7. The gas turbine end wall film hole layout structure based on the pit hole type according to claim 6, characterized in that: A simplified double-arc curved channel is obtained according to the pressure surface and suction surface profiles of the gas turbine turbine end wall, and the pressure gradient and secondary lateral flow characteristics of the simplified double-arc curved channel wall are kept similar to those of the cascade channel.

8. The gas turbine end wall film hole layout structure based on the pit hole type according to claim 7 is characterized in that: The turbine end wall of a gas turbine is the end wall of the first stage stationary blade or moving blade.

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