Turbine moving blade tip laminate cooling structure

Abstract

The invention provides a blade tip laminate cooling structure of a turbine moving blade, and belongs to the technical field of gas turbine blade cooling, the turbine moving blade comprises a blade tip groove, a pressure side wall surface and a suction side wall surface, and the blade tip laminate cooling structure of the turbine moving blade comprises an air film cooling hole, an inner layer plate and an impact hole formed in the inner layer plate, the inner layer plate is connected between the pressure side wall face and the suction side wall face, the space between the pressure side wall face and the suction side wall face is divided into an upper cavity and a lower cavity through the inner layer plate, and the cavity between the inner layer plate and the bottom wall of the blade tip groove is a blade tip inner cavity. The air film cooling holes are located in the bottom wall of the blade tip groove and the pressure side wall face and the suction side wall face on the two sides of the blade tip inner cavity. According to the treatment scheme, the cooling effect of the turbine movable blade tip is improved, the blade tip base body temperature and the movable blade tip ablation risk are reduced, and it is guaranteed that the blade is safely used in high-temperature fuel gas, and the service life is long.

Description

Technical Field

[0001] The present application relates to the technical field of gas turbine turbine blade cooling, and in particular to a turbine blade tip layer plate cooling structure. Background Art

[0002] The development of modern gas turbines is pursuing higher power and thermal efficiency. The temperature of the gas before the turbine continues to increase, and the turbine blades are subjected to severe thermal load tests.

[0003] In the prior art, in order to prevent the turbine blades from contacting and rubbing against the stationary casing during rotation, a radial gap is provided between the blade tip and the casing, which is usually referred to as the tip gap. Since there is a pressure gradient between the pressure side and the suction side of the blade tip, part of the high-temperature combustion gas is accelerated through the tip gap driven by the pressure gradient, forming a tip leakage flow. The tip leakage flow not only causes serious tip leakage losses and reduces turbine efficiency, but also has a strong heating effect on the blade tip, causing high tip temperature and posing a great challenge to the material life.

[0004] The grooved blade tip structure is currently the most commonly used structure in turbine blade tips to reduce tip clearance leakage losses. However, the grooved blade tip structure will cause the high-temperature combustion gas leakage flow in the tip clearance to impact the bottom wall of the tip groove under the action of vortex, resulting in violent convective heat transfer on the wall surface, and then generate higher temperatures and heat loads on the wall surface in the tip area, making the high-temperature area of ​​the blade tip prone to ablation, resulting in a reduced blade life, and thus threatening the reliability of the turbine.

[0005] In order to reduce the tip temperature, efficient cooling technology must be used. Tip cooling includes internal cooling and film cooling. In existing technologies, the internal cooling of the turbine blade tip generally adopts a cooling method of a radial straight channel or a serpentine channel from the blade root to the blade tip. However, these cooling methods have a small cooling gas flow rate and a slow flow rate near the blade tip, resulting in a low internal heat transfer coefficient and low tip cooling efficiency, which makes it difficult to meet the tip cooling requirements of turbine blades under high-temperature and high-pressure gas. For double-wall cooled blades with higher cooling efficiency, most of the cold air cools the walls of the blade body on both sides through the impact holes and film holes on the pressure and suction sides, and then merges into the mainstream gas. Even less cold air flows through the tip area, and the tip cooling problem of turbine blades under high-temperature and high-pressure gas has not yet been solved.

[0006] In view of the above shortcomings, it is necessary to design a turbine blade tip layer plate combined cooling structure to improve the tip cooling efficiency, reduce the blade tip base temperature, and ensure the safe use of turbine blades. Summary of the invention

[0007] In view of this, an embodiment of the present application provides a turbine blade tip layer plate cooling structure, which at least partially solves the problem in the prior art that the high-temperature combustion gas leakage flow at the blade tip impacts the bottom wall of the tip groove under the influence of the vortex, forming a high heat exchange area on the tip wall, making the high-temperature area of ​​the blade tip easy to burn, thereby proposing a turbine blade tip layer plate cooling structure to reduce the tip base temperature and ensure the safe use of the turbine blade.

[0008] An embodiment of the present application provides a turbine blade tip layer plate cooling structure, the turbine blade comprising a tip groove, a pressure side wall surface and a suction side wall surface, the turbine blade tip layer plate cooling structure comprising an air film cooling hole, an inner layer plate and an impact hole arranged on the inner layer plate, the inner layer plate is connected between the pressure side wall surface and the suction side wall surface, the space between the pressure side wall surface and the suction side wall surface is divided into two upper and lower chambers by the inner layer plate, the chamber between the inner layer plate and the bottom wall of the blade tip groove is a blade tip inner cavity, and the air film cooling hole is located on the bottom wall of the blade tip groove and the pressure side wall surface and the suction side wall surface on both sides of the blade tip inner cavity.

[0009] According to a specific implementation of the embodiment of the present application, the film cooling holes on the bottom wall of the blade tip groove and the impact holes on the inner plate are staggered in position.

[0010] According to a specific implementation method of an embodiment of the present application, it also includes at least one short baffle, which is arranged in the inner cavity of the blade tip, one end of the short baffle is connected to the pressure side wall surface, and the other end of the short baffle is connected to the suction side wall surface, and the inner cavity of the blade tip is divided into multiple chambers by the short baffle.

[0011] According to a specific implementation method of an embodiment of the present application, it also includes a plurality of short baffles, which are arranged in the inner cavity of the blade tip, one end of the short baffle is connected to the pressure side wall or the suction side wall, and the other end of the short baffle is a free end. The short baffles connected to the pressure side wall and the short baffles connected to the suction side wall are arranged at intervals, and a serpentine rotating microchannel is formed by the plurality of short baffles.

[0012] According to a specific implementation of the embodiment of the present application, at least one of the upper side of the bottom wall of the blade tip groove and the upper side of the inner plate is provided with a spoiler structure.

[0013] According to a specific implementation of the embodiment of the present application, the spoiler structure is configured as a spoiler rib or a spoiler column.

[0014] According to a specific implementation method of an embodiment of the present application, the film cooling holes located on the pressure side wall and the suction side wall are arranged obliquely, and the inclination direction of the film cooling holes is along the outlet end of the gas flow path toward one side of the bottom wall of the blade tip groove.

[0015] According to a specific implementation of the embodiment of the present application, the equivalent diameter of the impact hole is set to 0.5 mm~2.0 mm.

[0016] According to a specific implementation of the embodiment of the present application, the ratio of the impact distance between the impact hole and the bottom wall of the blade tip groove to the equivalent diameter of the impact hole is 2 to 5.

[0017] According to a specific implementation of the embodiment of the present application, the ratio of the flow area of ​​the impact hole to the flow area of ​​the film cooling hole is 1 to 4.

[0018] Beneficial effects: The turbine blade tip layer plate cooling structure in the embodiment of the present application forms a blade tip inner cavity by setting an inner layer plate, opening a plurality of impact holes on the inner layer plate, and opening a plurality of air film holes on the pressure side wall and the suction side wall. On the one hand, the inner layer plate allows the cold air to impact and cool the bottom wall of the blade tip groove, strengthen the cold air flow in the blade tip inner cavity, and enhance the cooling effect of the blade tip inner cavity. On the other hand, the plurality of impact holes and air film cooling holes can increase the cold air flow of the blade tip air film hole, and enhance the flow in the blade tip inner cavity and the air film cooling effect. At the same time, the increase in the cold air flow of the air film cooling hole can reduce the tip gap leakage flow rate and flow velocity, thereby reducing the tip gap leakage loss. The use of this structure greatly improves the cooling effect of the turbine blade tip, reduces the tip base temperature, reduces the risk of blade tip ablation, and ensures the safe and long-life use of the blade in high-temperature combustion gas. BRIEF DESCRIPTION OF THE DRAWINGS

[0019] In order to more clearly illustrate the technical solutions of the embodiments of the present application, the drawings required for use in the embodiments will be briefly introduced below. Obviously, the drawings described below are only some embodiments of the present application. For ordinary technicians in this field, other drawings can be obtained based on these drawings without paying creative work.

[0020] Figure 1 A schematic diagram of the positional relationship between a moving blade and a casing according to an embodiment of the present invention; Figure 2 for Figure 1 AA side cross-sectional view in; Figure 3 for Figure 1 BB side cross-sectional view in.

[0021] In the figure: 1. turbine moving blade; 2. blade tip groove; 3. casing; 11. high-temperature combustion gas; 12. blade tip leakage flow; 13. cooling gas; 20. pressure side wall; 21. blade tip groove bottom wall; 22. suction side wall; 23. inner plate; 24. impact hole; 25. film cooling hole; 26. blade tip cavity; 27. short partition; 28. spoiler rib; 29. ​​spoiler column; 30. blade tip I cavity; 31; blade tip II cavity; 32. blade tip III cavity; 33. blade tip IV cavity. DETAILED DESCRIPTION

[0022] The embodiments of the present application are described in detail below with reference to the accompanying drawings.

[0023] The following describes the implementation methods of the present application through specific examples, and those skilled in the art can easily understand other advantages and effects of the present application from the contents disclosed in this specification. Obviously, the described embodiments are only a part of the embodiments of the present application, rather than all the embodiments. The present application can also be implemented or applied through other different specific implementation methods, and the details in this specification can also be modified or changed in various ways based on different viewpoints and applications without departing from the spirit of the present application. It should be noted that, in the absence of conflict, the following embodiments and the features in the embodiments can be combined with each other. Based on the embodiments in the present application, all other embodiments obtained by ordinary technicians in the field without making creative work belong to the scope of protection of the present application.

[0024] It should be noted that various aspects of the embodiments within the scope of the appended claims are described below. It should be apparent that the aspects described herein can be embodied in a wide variety of forms, and any specific structure and / or function described herein is merely illustrative. Based on the present application, it should be understood by those skilled in the art that an aspect described herein can be implemented independently of any other aspect, and two or more of these aspects can be combined in various ways. For example, any number of aspects described herein can be used to implement the device and / or practice the method. In addition, other structures and / or functionalities other than one or more of the aspects described herein can be used to implement this device and / or practice this method.

[0025] It should also be noted that the illustrations provided in the following embodiments are only schematic illustrations of the basic concept of the present application. The illustrations only show the components related to the present application rather than being drawn according to the number, shape and size of the components in actual implementation. In actual implementation, the type, quantity and proportion of each component may be changed at will, and the component layout may also be more complicated.

[0026] Additionally, in the following description, specific details are provided to facilitate a thorough understanding of the examples. However, it will be understood by those skilled in the art that the aspects described may be practiced without these specific details.

[0027] The present application embodiment provides a turbine blade tip plate cooling structure. Figures 1 to 3 Describe in detail.

[0028] In one embodiment, a turbine blade tip layer plate cooling structure is provided, wherein the turbine blade 1 includes a tip groove 2, a pressure side wall 20 and a suction side wall 22, the turbine blade tip layer plate cooling structure includes a film cooling hole 25, an inner layer plate 23 and an impact hole 24 arranged on the inner layer plate 23, the inner layer plate 23 is connected between the pressure side wall 20 and the suction side wall 22, the space between the pressure side wall 20 and the suction side wall 22 is divided into two upper and lower chambers by the inner layer plate 23, the chamber between the inner layer plate 23 and the bottom wall 21 of the blade tip groove is a blade tip inner cavity 26, and the film cooling hole 25 is located on the pressure side wall 20 and the suction side wall 22 on both sides of the blade tip groove bottom wall 21 and the blade tip inner cavity 26.

[0029] In specific implementation, the blade tip groove is located on the inner side of the casing 3, and a blade tip leakage flow 12 is formed in the blade tip groove. By setting the inner plate 23, the space between the pressure side wall surface 20 and the suction side wall surface 22 is divided into two upper and lower chambers. The chamber between the inner plate 23 and the bottom wall 21 of the blade tip groove is the blade tip inner cavity 26, which can enhance the cooling effect of the blade tip inner cavity 26 by strengthening the flow of cooling gas 13 in the blade tip inner cavity 26. In this embodiment, this effect is achieved by setting the impact hole 24 on the inner plate 23. Specifically, the cooling gas 13 enters the blade tip inner cavity 26 through the impact hole 24 on the inner plate 23, and the porous impact heat exchange takes away a large amount of heat from the bottom wall 21 of the blade tip groove, and finally flows out through a large number of air film cooling holes 25, forming a cooling air film to protect the side wall surface of the high-temperature combustion gas 11 of the blade, thereby enhancing the cooling effect of the blade in many aspects and greatly reducing the temperature of the base of the moving blade tip.

[0030] In this embodiment, the inner plate 23 is provided to form the blade tip cavity 26, a plurality of impact holes 24 are provided on the inner plate 23, and a plurality of film cooling holes 25 are provided on the pressure side wall 20 and the suction side wall 22. On the one hand, the inner plate 23 allows the cold air to impact and cool the bottom wall 21 of the blade tip groove, strengthen the cold air flow in the blade tip cavity 26, and enhance the cooling effect of the blade tip cavity 26. On the other hand, the plurality of impact holes 24 and film cooling holes 25 can increase the cold air flow of the blade tip film cooling hole 25, and enhance the flow and film cooling effect of the blade tip cavity 26. At the same time, the increase in the cold air flow of the film cooling hole 25 can reduce the flow and flow velocity of the blade tip leakage flow 12, and reduce the tip clearance leakage loss. The use of this structure greatly improves the cooling effect of the blade tip of the turbine moving blade 1, reduces the tip base temperature, reduces the risk of blade tip ablation, and ensures the safe and long-life use of the blade in the high-temperature combustion gas 11.

[0031] In one embodiment, the film cooling holes 25 on the blade tip groove bottom wall 21 are staggered with the impact holes 24 on the inner plate 23. The staggered hole positions further enhance the flow of the cooling gas 13 and the cooling effect.

[0032] In one embodiment, at least one short baffle 27 is further included, and the short baffle 27 is arranged in the blade tip inner cavity 26. One end of the short baffle 27 is connected to the pressure side wall 20, and the other end of the short baffle 27 is connected to the suction side wall 22. The blade tip inner cavity 26 is divided into a plurality of chambers by the short baffle 27.

[0033] In specific implementation, a short partition 27 for connecting the bottom wall 21 of the blade tip groove and the inner plate 23 is provided in the blade tip inner cavity 26, that is, the upper side of the short partition 27 is connected to the bottom wall 21 of the blade tip groove, the lower side is connected to the inner plate 23, and the left and right sides are respectively connected to the pressure side wall surface 20 and the suction side wall surface 22. Therefore, by using the short partition 27, the blade tip inner cavity 26 is divided into four chambers, which are the blade tip I cavity 30, the blade tip II cavity 31, the blade tip III cavity 32 and the blade tip IV cavity 33 from the leading edge to the trailing edge. The blade tip inner cavity 26 is divided into chambers to better control the flow distribution of each chamber. At the same time, the connection between the bottom wall 21 of the blade tip groove and the inner plate 23 can reduce the temperature difference between the side wall surface of the blade high-temperature combustion gas 11 and the inner plate 23, and reduce the thermal stress caused by the temperature difference.

[0034] In another embodiment, it also includes a plurality of short baffles 27, wherein the short baffles 27 are arranged in the blade tip inner cavity 26, one end of the short baffle 27 is connected to the pressure side wall 20 or the suction side wall 22, and the other end of the short baffle 27 is a free end, and the short baffles 27 connected to the pressure side wall 20 and the short baffles 27 connected to the suction side wall 22 are arranged at intervals, and a serpentine rotating microchannel is formed by the plurality of short baffles 27.

[0035] During specific implementation, a short partition 27 is provided in the blade tip inner cavity 26 to form a rotary microchannel. The cooling gas 13 impacts the bottom wall 21 of the blade tip groove through the impact hole 24, then flows through the rotary microchannel, and finally flows out from the film cooling hole 25 and the tail gap, which greatly enhances the convection capacity of the cooling gas 13 inside the blade tip inner cavity 26, strengthens the cooling effect of the blade tip inner cavity 26, and reduces the base temperature of the moving blade tip.

[0036] Preferably, the short baffle 27 for dividing the blade tip cavity 26 into a plurality of chambers and the short baffle 27 for forming a serpentine rotating microchannel can be provided at the same time. For example, a short baffle 27 rotating microchannel can be provided in the blade tip IV cavity 33, which can effectively control the flow distribution of each chamber, further enhance the convection capacity of the cooling gas 13 inside the blade tip cavity 26, strengthen the cooling effect of the blade tip cavity 26, and reduce the temperature of the base of the moving blade tip. It should be noted that, according to needs, a short baffle 27 for forming a serpentine rotating microchannel can also be provided in other chambers.

[0037] In one embodiment, at least one of the upper side of the blade tip groove bottom wall 21 and the upper side of the inner layer plate 23 is provided with a spoiler structure.

[0038] Furthermore, the spoiler structure is configured as a spoiler rib 28 or a spoiler column 29 .

[0039] In specific implementation, the spoiler column 29 or the spoiler rib 28 can be arranged on the bottom wall 21 of the blade tip groove, or on the inner plate 23, such as Figure 3 As shown, this embodiment only shows the case where it is set on the inner plate 23, and the spoiler structure can also be set on the tip groove bottom wall 21 according to the situation. According to the implementation situation, the spoiler column 29 and the spoiler rib 28 can also be set at the same time to enhance the spoiler effect.

[0040] Specifically, the spoiler rib 28 can be set to a herringbone spoiler rib 28 structure, such as Figure 3 As shown, a herringbone spoiler rib 28 structure is provided in the blade tip III cavity 32. The cooling gas 13 impacts the bottom wall 21 of the blade tip groove through the impact hole 24 and then flows through the spoiler rib 28, and is finally discharged from the pressure side air film cooling hole 25, thereby enhancing the convection capacity of the cooling gas 13 inside the blade tip inner cavity 26, strengthening the cooling effect of the blade tip inner cavity 26, and reducing the base temperature of the moving blade tip.

[0041] In one embodiment, the film cooling holes 25 located on the pressure side wall 20 and the suction side wall 22 are inclined, and the inclination direction of the film cooling holes 25 is along the outlet end of the gas flow path toward one side of the tip groove bottom wall 21.

[0042] In one embodiment, the equivalent diameter of the impact hole 24 is set to 0.5 mm to 2.0 mm. Under the same cold air flow condition, the smaller the impact hole 24 diameter is, the larger the impact Reynolds number is, and the better the average heat exchange effect of the impact target surface is. The appropriate impact hole 24 diameter can be selected in combination with the actual processing technology.

[0043] In one embodiment, the ratio of the impact distance between the impact hole 24 and the bottom wall 21 of the blade tip groove to the equivalent diameter of the impact hole 24 is 2 to 5. If the impact distance is too small, the jet is not fully developed, and the impact jet contacts the wall just after leaving the layer plate, resulting in insufficient heat exchange. If the impact distance is too large, the energy loss of the jet along the way increases, the flow velocity on the wall is reduced, and the impact cooling effect of the target surface is also reduced. Within this setting range, the impact jet is fully developed, the flow velocity of the lateral flow is relatively low, the loss of the impact jet passing through the lateral flow is small, and the comprehensive cooling effect of the impact target surface is optimal.

[0044] In one embodiment, the ratio of the flow area of ​​the impact hole 24 to the flow area of ​​the film cooling hole 25 is 1 to 4. This numerical range takes into account both the cooling effect and the cooling reliability. If the flow area of ​​the impact hole 24 is too small, it may lead to a small cold air flow, a low pressure in the blade tip cavity 26, a backflow of the film cooling hole 25, and the gas invading the interior of the blade from the film cooling hole 25. If the flow area of ​​the impact hole 24 is too large, it may lead to a large cold air flow, a high pressure in the blade tip cavity 26, and a high flow rate of the cold air jet flowing out of the film cooling hole 25, so that a cooling film cannot be formed on the wall, resulting in a significant reduction in the overall cooling effect.

[0045] In the embodiment provided by the present invention, an inner plate 23 is provided to form a blade tip cavity 26, a plurality of impact holes 24 are provided on the inner plate 23, and a plurality of film cooling holes 25 are provided on the pressure side wall 20 and the suction side wall 22. On the one hand, the inner plate 23 allows cold air to impact and cool the bottom wall 21 of the blade tip groove, and strengthens the cold air flow in the blade tip cavity 26 through the rotating microchannel, thereby enhancing the cooling effect of the blade tip cavity 26. On the other hand, the plurality of impact holes 24 and film cooling holes 25 can increase the cold air flow rate of the blade tip film cooling hole 25, thereby enhancing the flow and film cooling effect of the blade tip cavity 26. At the same time, the increase in the cold air flow rate of the film cooling hole 25 can reduce the tip gap leakage flow rate and flow velocity, thereby reducing the tip gap leakage loss. The use of this structure greatly improves the cooling effect of the blade tip of the turbine moving blade 1, reduces the tip base temperature, reduces the risk of blade tip ablation, and ensures the safe and long-life use of the blade in the high-temperature combustion gas 11.

[0046] At the same time, an inner layer plate 23 is provided to form a blade tip cavity 26, and a plurality of short partitions 27 and a turbulent structure are provided in the blade tip cavity 26 to reduce the temperature difference between the side wall surface of the high-temperature combustion gas 11 of the blade and the inner layer plate 23, thereby reducing the thermal stress caused by the temperature difference of the blade and extending the service life of the moving blade.

[0047] The above is only a specific implementation of the present application, but the protection scope of the present application is not limited thereto. Any changes or substitutions that can be easily thought of by a person skilled in the art within the technical scope disclosed in the present application should be included in the protection scope of the present application. Therefore, the protection scope of the present application shall be based on the protection scope of the claims.

Claims

1. A turbine blade tip layer plate cooling structure, the turbine blade (1) comprising a blade tip groove (2), a pressure side wall surface (20) and a suction side wall surface (22), characterized in that: The turbine blade tip layer plate cooling structure comprises a film cooling hole (25), an inner layer plate (23) and an impact hole (24) arranged on the inner layer plate (23); the inner layer plate (23) is connected between the pressure side wall surface (20) and the suction side wall surface (22); the space between the pressure side wall surface (20) and the suction side wall surface (22) is divided into two upper and lower chambers by the inner layer plate (23); the chamber between the inner layer plate (23) and the bottom wall (21) of the blade tip groove is the blade tip inner cavity (26); the film cooling hole (25) is located on the pressure side wall surface (20) and the suction side wall surface (22) on both sides of the bottom wall (21) of the blade tip groove and the blade tip inner cavity (26).

2. The turbine blade tip plate cooling structure according to claim 1, characterized in that: The film cooling holes (25) on the bottom wall (21) of the blade tip groove and the impact holes (24) on the inner plate (23) are arranged at staggered positions.

3. The turbine blade tip plate cooling structure according to claim 1, characterized in that: The blade tip inner cavity (26) further comprises at least one short baffle (27), wherein the short baffle (27) is arranged in the blade tip inner cavity (26), one end of the short baffle (27) is connected to the pressure side wall surface (20), and the other end of the short baffle (27) is connected to the suction side wall surface (22), and the blade tip inner cavity (26) is divided into a plurality of chambers by the short baffle (27).

4. The turbine blade tip plate cooling structure according to claim 1, characterized in that: It also includes a plurality of short baffles (27), wherein the short baffles (27) are arranged in the blade tip inner cavity (26), one end of the short baffle (27) is connected to the pressure side wall surface (20) or the suction side wall surface (22), and the other end of the short baffle (27) is a free end. The short baffles (27) connected to the pressure side wall surface (20) and the short baffles (27) connected to the suction side wall surface (22) are arranged at intervals, and a serpentine rotating microchannel is formed by the plurality of short baffles (27).

5. The turbine blade tip plate cooling structure according to claim 1, characterized in that: At least one of the upper side of the blade tip groove bottom wall (21) and the upper side of the inner layer plate (23) is provided with a spoiler structure.

6. The turbine blade tip plate cooling structure according to claim 5, characterized in that: The spoiler structure is configured as a spoiler rib (28) or a spoiler column (29).

7. The turbine blade tip plate cooling structure according to claim 1, characterized in that: The air film cooling holes (25) located on the pressure side wall surface (20) and the suction side wall surface (22) are arranged at an inclination, and the inclination direction of the air film cooling holes (25) is along the outlet end of the gas flow path toward one side of the blade tip groove bottom wall (21).

8. The turbine blade tip plate cooling structure according to any one of claims 1 to 7, characterized in that: The equivalent diameter of the impact hole (24) is set to 0.5 mm to 2.0 mm.

9. The turbine blade tip plate cooling structure according to any one of claims 1 to 7, characterized in that: The ratio of the impact distance between the impact hole (24) and the blade tip groove bottom wall (21) to the equivalent diameter of the impact hole (24) is 2 to 5.

10. The turbine blade tip plate cooling structure according to any one of claims 1 to 7, characterized in that: The ratio of the flow area of ​​the impact hole (24) to the flow area of ​​the film cooling hole (25) is 1 to 4.

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