Turbine blade composite leading edge cooling structure integrating multi-stage enhanced heat transfer mechanisms

Abstract

This application provides a composite leading-edge cooling structure for turbine blades integrating a multi-stage enhanced heat transfer mechanism, comprising: a blade body; an impact plate dividing the channel into a cold air channel and a leading-edge cavity; a row of air inlets on the impact plate along the blade height direction; multiple film cooling holes located on and penetrating the leading-edge a-wall, arranged in multiple rows along the blade height direction; and multiple protruding structures arranged in multiple rows, with each protruding structure positioned between adjacent film cooling holes. This application achieves systematic thermal management of the high-heat-load region of the turbine blade leading edge through the synergistic coupling of three mechanisms: impact cooling, enhanced heat transfer through internal discrete protruding structures, and external staggered arrangement of adjacent rows of film cooling holes. This results in a more uniform temperature distribution on the leading-edge a-wall, enhanced stability and coverage of external film cooling, and effectively reduced internal flow resistance compared to traditional continuous rib structures.

Description

Technical Field

[0001] This application relates to the field of turbine blade technology, and in particular to a composite leading edge cooling structure for turbine blades that integrates a multi-stage enhanced heat transfer mechanism. Background Technology

[0002] As aero-engines evolve towards higher performance, turbine gas inlet temperatures continue to rise. Under the harsh environment of extreme high temperature and pressure, the thermal load on turbine blades increases dramatically. The blade leading edge, directly exposed to the hottest and highest-pressure incoming gas, becomes the critical area with the most concentrated heat load and is most susceptible to high-temperature oxidation, ablation, and thermal fatigue. Excessively high leading edge wall temperatures severely reduce material strength and lifespan, constituting a bottleneck restricting engine performance and reliability. Therefore, developing efficient and reliable blade leading edge cooling and thermal protection technologies has become an urgent need to overcome the thermal barrier, ensure safe engine operation, and extend the service life of critical components.

[0003] To improve heat exchange capacity, existing technologies often employ reinforced structures such as continuous ribs. While this can enhance heat exchange to some extent, it generally leads to increased flow resistance and a significant increase in pressure loss, which is detrimental to the overall efficiency of the cooling system and the improvement of engine performance. Summary of the Invention

[0004] In view of this, the purpose of this application is to provide a turbine blade composite leading edge cooling structure with integrated multi-stage enhanced heat transfer mechanism, so as to solve the problem of large flow loss caused by continuous ribs in the prior art.

[0005] Embodiments of this application provide a turbine blade composite leading-edge cooling structure integrating a multi-stage enhanced heat transfer mechanism, including: The blade body includes a leading edge a, and a channel is provided within the leading edge a; An impact plate is located within the channel along the blade height direction, and the impact plate divides the channel into a cold air channel and a leading edge cavity; a row of air inlets is provided on the impact plate along the blade height direction; Multiple film air pores are located on and penetrate the leading edge a wall. Along the blade height direction, the multiple film air pores are arranged in multiple rows, and the film air pores in adjacent rows are staggered. Multiple protrusions are located on the inner wall of the leading edge a along the blade height direction. The multiple protrusions are arranged in multiple rows, and each of the protrusions is arranged between adjacent air film pores.

[0006] In some embodiments, a row of air inlets and a middle row of air film inlets are arranged alternately.

[0007] In some embodiments, each of the protrusion structures is one or more of an elliptic convex sphere, a hemispherical convex sphere, and an oblate spheroid.

[0008] In some embodiments, the number of rows of protrusions is the same as the number of rows of air film pores.

[0009] In some embodiments, a row of air holes is located on the centerline of the impact plate along the blade height direction.

[0010] In some embodiments, the pore size of each of the air film pores is d The hole is 5 cm long. d The distance between adjacent air film pores is 5 d , d The thickness is 0.4mm to 0.8mm.

[0011] In some embodiments, the angle between the axial direction of the air film pore and the tangential direction of the blade body surface is . β The angle between the air film pore and the direction of blade height is α .

[0012] In some embodiments, 45°≤ β ≤90°; and / or 30°≤ α ≤60°.

[0013] In some embodiments, each of the protrusion structures is an elliptic convex sphere with a major axis of 3. d The minor axis of the elliptic convex sphere is 2. d The distance between adjacent elliptic convex spheres is 5. d , d The thickness is 0.4mm to 0.8mm.

[0014] In some embodiments, the diameter of the impact hole is 2. d The spacing between adjacent impact holes is 10. d , d The thickness is 0.4mm to 0.8mm.

[0015] Beneficial effects: This application constructs a complete synergistic system from internal flow modulation to external film cooling by synergistic coupling of three mechanisms in space and flow: impact cooling, internal discrete protrusion structure to enhance heat transfer through turbulence, and external film cooling with staggered arrangement of two adjacent rows of air film holes. This achieves systematic thermal management of the high-heat-load region of the turbine blade leading edge. Within a limited internal space, the structure of this application significantly improves the overall cooling efficiency of the leading edge a, making the temperature distribution of the leading edge a wall more uniform, enhancing the stability and coverage of external film cooling, and effectively reducing internal flow resistance compared to the traditional continuous rib structure. This application also effectively reduces the risk of thermal fatigue and oxidation at the leading edge of the blade body, providing a reliable technical solution for improving engine operating temperature and extending blade life.

[0016] The "multiple air film holes staggered along the blade height direction", "discrete protrusions staggered and corresponding to the air film holes along the blade height direction", and "directional guidance of the impinging jet to the protrusions" constitute a multi-stage synergistic cooling mechanism. This multi-stage synergistic cooling mechanism, while ensuring structural compactness, achieves systematic thermal management from the inside of the leading edge a to the outer surface. This effectively avoids the large flow loss caused by continuous ribs in the prior art, and solves the problem of internal and external synergistic enhancement not addressed in traditional solutions. Ultimately, it achieves a comprehensive improvement in cooling efficiency, optimized temperature uniformity, ensured air film coverage stability, and reduced flow resistance. It significantly improves the overall cooling effect and the uniformity of wall temperature distribution, effectively reduces the peak wall temperature, and helps extend the service life of the blade under high temperature and high pressure environments. Attached Figure Description

[0017] To more clearly illustrate the technical solutions of the embodiments of this application, the accompanying drawings of the embodiments will be briefly described below. Obviously, the drawings described below only relate to some embodiments of this application and are not intended to limit this application, wherein: Figure 1 This is a schematic diagram of the turbine blade composite leading edge cooling structure integrating a multi-stage enhanced heat transfer mechanism as described in this application. Figure 2 Left view of the turbine blade composite leading edge cooling structure integrating a multi-stage enhanced heat transfer mechanism as described in this application; Figure 3 For this application Figure 2 Sectional view of AA; Figure 4 This is a structural diagram of the turbine blade composite leading edge cooling structure integrating a multi-stage enhanced heat transfer mechanism, without an impact plate, as described in this application. Figure 5 For this application Figure 4 Enlarged view of B in the middle; Figure 6 This is a schematic diagram of the turbine blade composite leading edge cooling structure integrating a multi-stage enhanced heat transfer mechanism, as presented in this application, from another perspective.

[0018] In the figure: 1-blade body, 11-impact plate, 12-cooling air channel, 13-leading edge cavity, 14-film heating hole, 15-protruding structure, 110-air inlet. Detailed Implementation

[0019] The technical solutions of the embodiments of this application will be clearly and comprehensively described below with reference to the accompanying drawings. Obviously, the described embodiments are only some embodiments of this application, and not all embodiments. Based on the embodiments of this application, all other embodiments obtained by those skilled in the art without creative effort are within the scope of protection of this application.

[0020] In the description of this application, it should be understood that the terms "length," "width," "thickness," "upper," "lower," "vertical," "horizontal," "top," "bottom," "inner," and "outer," etc., indicating orientation or positional relationships based on the orientation or positional relationships shown in the accompanying drawings, are only for the convenience of describing this application and simplifying the description, and do not indicate or imply that the device or element referred to must have a unique orientation, or be constructed and operated in a specific orientation, and therefore should not be construed as a limitation of this application. Furthermore, the terms "first" and "second" are used for descriptive purposes only and should not be construed as indicating or implying relative importance or specifying the number of technical features indicated. Thus, features defined with "first" and "second" may explicitly or implicitly include one or more features. In the description of this application, "multiple" means two or more, unless otherwise explicitly specified.

[0021] In this application, the term "exemplary" is used to mean "serving as an example, illustration, or description." Any embodiment described as "exemplary" in this application is not necessarily to be construed as being more preferred or advantageous than other embodiments. The following description is provided to enable any person skilled in the art to make and use this application. Details are set forth in the following description for purposes of explanation. It should be understood that those skilled in the art will recognize that this application can be made without using these specific details. In other instances, known structures and processes are not described in detail to avoid obscuring the description of this application with unnecessary detail. Therefore, this application is not intended to be limited to the embodiments shown, but is consistent with the broadest scope of the principles disclosed in this application.

[0022] Please see Figure 1-3 and Figure 6 Embodiments of this application provide a turbine blade composite leading-edge cooling structure integrating a multi-stage enhanced heat transfer mechanism, comprising: The blade body 1 includes a leading edge a, and a channel is provided in the leading edge a; An impact plate 11 is located within the channel along the blade height direction, and the impact plate 11 divides the channel into a cold air channel 12 and a leading edge cavity 13; a row of air inlets 110 is provided on the impact plate 11 along the blade height direction; Multiple film air holes 14 are located on and penetrate the leading edge a wall. Along the blade height direction, the multiple film air holes 14 are arranged in multiple rows, and the film air holes 14 in adjacent rows are staggered. Multiple protrusions 15 are located on the inner wall of the leading edge a along the blade height direction. The multiple protrusions 15 are arranged in multiple rows, and each of the protrusions 15 is arranged between adjacent air film holes 14.

[0023] During operation, cooling gas is drawn from the engine compressor and delivered to the root of the leading edge a of the blade body 1 via the cooling gas passage 12. When the airflow passes through the impact holes on the impact plate 11, the flow area decreases abruptly, forming multiple high-speed impact jets. After these jets enter the leading edge cavity 13, some directly impact the surface of the protruding structure 15. The impact process not only greatly enhances the local heat transfer in the area of ​​the protruding structure 15, but the high-speed airflow also deflects and breaks up after impacting the protruding structure 15, inducing complex swirling and secondary flows within the leading edge cavity 13. These swirling flows further scour other areas of the inner wall of the leading edge a, destroying the low-speed stagnation zone and thermal boundary layer near the entire wall surface, achieving efficient and uniform convective cooling of the entire inner wall surface of the leading edge a.

[0024] The cooling gas, after being impacted and swirled for heat exchange inside the leading edge cavity 13, experiences a decrease in kinetic energy during repeated interactions with the leading edge a wall and the protruding structure 15, resulting in a smoother flow velocity. Subsequently, the airflow exits through staggered film gas holes 14 on the outer wall of the leading edge a. Because the internal flow has been sufficiently disturbed and homogenized, uniform and stable inflow conditions are provided for the film gas holes 14, ensuring that the velocity of the cold air discharged from each film gas hole 14 is moderate and consistent. The staggered arrangement of the film gas holes 14 allows the cooling film formed by the preceding row of film jets to spatially intersect and complement the film formed by the following row of jets, thereby forming a continuous, complete, and uniformly covered film gas protective layer on the outer surface of the leading edge a, effectively isolating it from the direct impact of external high-temperature combustion gases.

[0025] It is understandable that the multiple exhaust film holes 14 are located on the leading edge a wall where the leading edge cavity 13 is located.

[0026] It is understood that multiple air film holes 14 can be provided as needed. For example, three air film holes 14 are provided, as well as three rows of protruding structures 15, and each protruding structure 15 is arranged between adjacent air film holes 14.

[0027] It is understandable that after the cold air impact jet enters the inner cavity of the leading edge, it is discharged through the gas film hole 14 and forms a continuous cooling gas film on the outer surface of the leading edge, thereby effectively isolating the high-temperature combustion gas.

[0028] It is understandable that the advantage of the staggered arrangement of the air film holes 14 in the two adjacent rows is that the cooling barrier formed by the front air film jet and the cooling barrier formed by the rear air film jet complement each other in space, which significantly improves the continuity and uniformity of the air film coverage and effectively avoids the formation of local high heat flux areas due to insufficient jet lifting or diffusion.

[0029] It is understandable that after the cooling gas enters from the cold air channel 12, it forms a high-speed jet through the impact hole and enters the leading edge cavity 13.

[0030] It is understood that the "along the leaf height direction" in this application refers to the height direction along the blade body 1.

[0031] It is understandable that the purpose of setting up multiple protrusions 15 is to actively disturb the cooling airflow near the inner wall of the leading edge, destroy the thermal boundary layer, enhance local convective heat transfer, and expand the heat transfer area.

[0032] It is understood that each of the protruding structures 15 is arranged between adjacent air film holes 14 and forms a spatial coupling with multiple rows of staggered air film holes 14. Its advantage is that the local enhanced heat transfer area generated by the protruding structure 15 is located precisely in the near-wall transverse gap area formed by the staggered multiple air film holes 14 where the air film coverage is relatively weak. This not only compensates for the heat transfer weakness in this area, but also provides a more uniform and stable inflow to the downstream staggered multiple air film holes 14 by enhancing the cold air disturbance. Thus, it achieves precise complementarity and synergistic enhancement of internal convection enhancement and external air film coverage in space.

[0033] In this application, the flow area of ​​the cold airflow decreases sharply when passing through the impact hole, forming a high-speed impact jet. After entering the leading edge cavity 13, part of the airflow forms a swirling flow and interacts with the discretely arranged protruding structures 15 to generate local vortices, effectively enhancing convective heat transfer. This design significantly enhances disturbance and heat transfer intensity while avoiding the large-scale flow resistance and pressure loss caused by continuous ribs. It achieves efficient synergy between impact cooling, internal swirling enhancement, and external film cooling, systematically improving the overall cooling efficiency of the leading edge.

[0034] In this application, the impinging jet is precisely guided to the region of the raised structure 15. Combined with the directional disturbance of the wall boundary layer by the discrete raised structure 15, this effectively eliminates the low-velocity stagnation zone and localized hot spots that are prone to occur in traditional structures. The raised structure 15 is arranged along the blade height direction corresponding to the weak heat transfer zone between the rows of film air holes 14, compensating for insufficient external film coverage, resulting in a more uniform heat load distribution on the leading edge a-wall, and improving the reliability and durability of thermal protection.

[0035] When the internal swirling flow passes through the region of the protruding structure 15, the momentum is dissipated and modulated in an orderly manner, resulting in a smoother flow velocity and significantly enhanced adhesion of the cold air discharged from the film cooling holes 14. Combined with the spatial complementary cooling barrier formed by multiple rows of staggered film cooling holes 14, the stability, continuity, and uniformity of the film cooling layer on the outer surface of the leading edge a are significantly improved, thereby effectively isolating high-temperature combustion gases and enhancing resistance to mainstream stripping.

[0036] Please see Figure 1 In some embodiments, a row of air inlets 110 and a middle row of air film inlets 14 are staggered.

[0037] It is understood that the row of air inlets 110 and the middle row of film cooling holes 14 are staggered, allowing the impact jet to be precisely guided to the area of ​​the protruding structure 15. On the one hand, the jet directly impacts the surface of the protruding structure 15, enhancing local disturbance and heat transfer; on the other hand, the reduced airflow velocity and momentum after the impact help to promote the subsequent extension of the cool air discharged from the film cooling holes 14 outward at a lower speed and with higher adhesion, thereby forming a more stable and more uniformly covered cooling film on the outer surface of the blade leading edge.

[0038] For example, the position of the impact hole is offset from the exhaust film hole 14 at the center line of the leading edge a. This design allows the high-speed airflow ejected from the impact hole to directly impact the area of ​​the protruding structure 15 arranged on the inner wall of the leading edge a.

[0039] Please see Figure 3 and Figure 5 In some embodiments, each of the protrusion structures 15 is one or more of an elliptical convex sphere, a hemispherical convex sphere, and an oblate spheroid.

[0040] Please see Figure 4 Furthermore, the number of rows of protruding structures 15 is the same as the number of rows of air film holes 14.

[0041] Please see Figure 1 In some embodiments, along the blade height direction, a row of the air holes 110 is located on the center line of the impact plate 11.

[0042] For example, the impact plate 11 is positioned directly opposite the multi-vent membrane hole 14, which is located on the leading edge a wall where the leading edge cavity 13 is located.

[0043] It is understandable that a row of air holes 110 is located on the center line of the impact plate 11 along the blade height direction. The purpose is to ensure that when the cold air impacts the inner surface of the leading edge, it can spread evenly to both sides of the center line of the leading edge.

[0044] In some embodiments, the diameter of each air film hole 14 is d, the length of the hole is 5d, and the distance between adjacent air film holes 14 is 5d, where d is 0.4mm to 0.8mm, for example, it can be 0.4mm, 0.45mm, 0.5mm, 0.55mm, 0.6mm, 0.65mm, 0.7mm, 0.75mm, 0.8mm, etc.

[0045] In some embodiments, the angle between the axial direction of the air film pore 14 and the tangential direction of the blade body 1 is β, and the angle between the channel of the air film pore 14 and the blade height direction is α.

[0046] Furthermore, 45°≤β≤90°, where β can be, for example, 45°, 48°, 50°, 52°, 55°, 56°, 58°, 60°, 62°, 65°, 68°, 70°, 72°, 75°, 78°, 80°, 82°, 85°, 88°, 90°, etc.

[0047] Furthermore, 30°≤α≤60°, where α can be, for example, 30°, 32°, 35°, 38°, 40°, 42°, 45°, 48°, 50°, 52°, 55°, 58°, 60°, etc.

[0048] In some embodiments, each of the protrusion structures 15 is an elliptical convex sphere with a major axis of 3d, a minor axis of 2d, and a distance of 5d between adjacent elliptical convex spheres. d is 0.4mm to 0.8mm, for example, 0.4mm, 0.45mm, 0.5mm, 0.55mm, 0.6mm, 0.65mm, 0.7mm, 0.75mm, 0.8mm, etc.

[0049] In some embodiments, the diameter of the impact hole is 2d, and the spacing between adjacent impact holes is 10d, where d is 0.4mm to 0.8mm, for example, it can be 0.4mm, 0.45mm, 0.5mm, 0.55mm, 0.6mm, 0.65mm, 0.7mm, 0.75mm, 0.8mm, etc.

[0050] The basic concepts have been described above. Obviously, for those skilled in the art, the detailed disclosure above is merely illustrative and does not constitute a limitation of this application. Although not explicitly stated herein, those skilled in the art may make various modifications, improvements, and corrections to this application. Such modifications, improvements, and corrections are suggested in this application, and therefore remain within the spirit and scope of the exemplary embodiments of this application.

[0051] Furthermore, this application uses specific terms to describe embodiments of the application. For example, "an embodiment," "one embodiment," and / or "some embodiments" refer to a particular feature, structure, or characteristic associated with at least one embodiment of the application. Therefore, it should be emphasized and noted that "an embodiment," "one embodiment," or "an alternative embodiment" mentioned twice or more in different locations in this specification do not necessarily refer to the same embodiment. In addition, certain features, structures, or characteristics in one or more embodiments of the application can be appropriately combined.

Claims

1. A turbine blade composite leading edge cooling structure integrating a multi-stage enhanced heat transfer mechanism, characterized in that, include: The blade body includes a leading edge a, and a channel is provided within the leading edge a; An impact plate is located within the channel along the blade height direction, and the impact plate divides the channel into a cold air channel and a leading edge cavity; a row of air inlets is provided on the impact plate along the blade height direction; Multiple film air pores are located on and penetrate the leading edge a wall. Along the blade height direction, the multiple film air pores are arranged in multiple rows, and the film air pores in adjacent rows are staggered. Multiple protrusions are located on the inner wall of the leading edge a along the blade height direction. The multiple protrusions are arranged in multiple rows, and each of the protrusions is arranged between adjacent air film pores.

2. The turbine blade composite leading edge cooling structure with integrated multi-stage enhanced heat transfer mechanism according to claim 1, characterized in that, The row of air inlets and the middle row of air film inlets are staggered.

3. The turbine blade composite leading edge cooling structure with integrated multi-stage enhanced heat transfer mechanism according to claim 1, characterized in that, Each of the aforementioned protrusions is one or more of the following: an elliptical convex sphere, a hemispherical convex sphere, and an oblate spheroid.

4. The turbine blade composite leading edge cooling structure with integrated multi-stage enhanced heat transfer mechanism according to claim 3, characterized in that, The number of rows of raised structures is the same as the number of rows of air film pores.

5. The turbine blade composite leading edge cooling structure with integrated multi-stage enhanced heat transfer mechanism according to claim 1, characterized in that, Along the blade height direction, a row of air holes is located on the center line of the impact plate.

6. The turbine blade composite leading edge cooling structure with integrated multi-stage enhanced heat transfer mechanism according to claim 1, characterized in that, The diameter of each of the air film pores is d The hole is 5 cm long. d The distance between adjacent air film pores is 5 d , d The thickness is 0.4mm to 0.8mm.

7. The turbine blade composite leading edge cooling structure with integrated multi-stage enhanced heat transfer mechanism according to claim 1, characterized in that, The angle between the axis of the film pore and the tangential direction of the blade body surface is... β The angle between the air film pore and the direction of blade height is α .

8. The turbine blade composite leading edge cooling structure with integrated multi-stage enhanced heat transfer mechanism according to claim 7, characterized in that, 45°≤ β ≤90°; and / or 30°≤ α ≤60°。 9. The turbine blade composite leading edge cooling structure with integrated multi-stage enhanced heat transfer mechanism according to claim 1, characterized in that, Each of the aforementioned protrusions is an elliptic convex sphere, with a major axis of 3. d The minor axis of the elliptic convex sphere is 2. d The distance between adjacent elliptic convex spheres is 5. d , d The thickness is 0.4mm to 0.8mm.

10. The turbine blade composite leading edge cooling structure with integrated multi-stage enhanced heat transfer mechanism according to claim 1, characterized in that, The diameter of the impact hole is 2 d The spacing between adjacent impact holes is 10. d , d The thickness is 0.4mm to 0.8mm.

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