Turbine blade trailing edge cooling structure, design method and aero-engine

Abstract

The invention discloses a turbine blade tail edge cooling structure, a design method and an aero-engine, the tail edge of a blade is provided with a first cavity communicated to an inner cavity of the blade and a second cavity communicated to the exhaust end of the tail edge of the blade and the first cavity, and an impact air entraining structure is arranged between the first cavity and the second cavity. The inner wall of the second cavity is provided with turbulent flow structures, the turbulent flow structures are boss structures formed on the inner wall of the second cavity, the radial arrangement positions of the turbulent flow structures are matched with the radial arrangement positions of the impact air entraining structures, and the turbulent flow structures on the inner wall of the blade basin side and the turbulent flow structures on the inner wall of the blade back side of the second cavity are arranged in a staggered mode. Compared with a turbulent flow column array structural form, the structural form effectively enhances heat exchange, impact air entraining enables high-speed cold air to impact the surface of an inner cavity of the blade, high-strength convective heat exchange is generated at the position needing key cooling, and the local high temperature is effectively reduced. And the boss can cause disturbance in the boundary layer, so that heat exchange between the cold air and the wall surface is more sufficient, and the local heat exchange effect is further enhanced.

Description

Technical Field

[0001] The present invention relates to the technical field of turbine blades, and in particular, to a trailing edge cooling structure of a turbine blade. In addition, the present invention also relates to an aeroengine including the above-mentioned trailing edge cooling structure of the turbine blade. Background Art

[0002] In the process of continuous innovation in high-performance aeroengine and gas turbine technologies, the thermodynamic cycle parameters of the engine have continuously broken through the upper limit. This development trend has made the thermal protection problem of turbine blades, which are key hot-end components, increasingly difficult. Due to the fact that the blades are in a harsh environment of high temperature, high pressure and high-speed gas flow scouring for a long time, the requirements for their thermal protection measures are extremely strict. The air-cooling technology is the main means to solve the thermal protection problem of turbine blades. By carefully planning the cold air flow path inside the blade, an extremely complex cold air flow path is constructed to ensure the stable operation of the blade under high-temperature conditions.

[0003] The basic cooling methods of aeroengine turbine blades mainly include convective cooling, impingement cooling and film cooling. In practical applications, in order to achieve the best cooling effect, most cooling blades adopt a composite cooling design, combining one or more of the above cooling methods. Especially in the trailing edge part of the turbine blade, a turbulator array is often set. These turbulators connect the blade back and the blade basin in the form of one row or multiple rows, and their effects are significant. On the one hand, the turbulators can cause strong convection of the cold air in the turbulator area, greatly improving the heat transfer capacity. On the other hand, they can enhance the structural strength of the blade in the trailing edge area, making it sufficient to withstand the ultra-high aerodynamic load on the gas side. After a large number of practices and optimizations, the turbulator array enhanced cooling structure has been very mature, and this structure has been widely used in many engines under research and in service at present.

[0004] The main disadvantages of the above technical solutions are reflected in the following aspects: 1) Under the conditions of limited space and limited cold air volume, the cooling effect of the turbulator array cooling structure has an obvious upper limit value, and it is difficult to improve the cooling effect of the trailing edge area by simply changing the turbulator parameters; 2) The cold air flow resistance in the single turbulator array area is very large, which will cause complex vortices and wakes of the air flow around the columns, increasing the degree of flow disorder and resistance, and requiring a high cold air supply pressure to ensure sufficient cold air volume, and the cost of realizing this in the engine is very high; 3) The shape structure and arrangement method are complex, which are easily restricted by the narrow space at the trailing edge of the blade, and have high requirements for manufacturing processes, die accuracy, etc. Summary of the Invention

[0005] The present invention provides a trailing edge cooling structure for a turbine blade, a design method and an aero-engine, so as to solve the technical problems in the prior art that it is difficult to improve the cooling effect of the turbulator array structure form, the flow resistance is large, resulting in a large demand for cooling air volume and poor stability, and the structure is complex, resulting in poor adaptability and high processing difficulty.

[0006] According to one aspect of the present invention, there is provided a trailing edge cooling structure for a turbine blade. The trailing edge of the blade has a first chamber communicating with the inner cavity of the blade, and a second chamber communicating with the exhaust end and the first chamber of the trailing edge of the blade respectively. An impact air intake structure is arranged between the first chamber and the second chamber, and a flow disturbing structure is arranged on the inner wall of the second chamber. The flow disturbing structure is a boss structure formed on the inner wall of the second chamber. The radial layout position of the flow disturbing structure matches the radial layout position of the impact air intake structure. The flow disturbing structures on the inner wall of the blade's suction side and the inner wall of the blade's pressure side in the second chamber are arranged alternately.

[0007] As a further improvement of the above technical solution, the impact air intake structure includes a plurality of impact holes uniformly distributed along the radial direction of the blade.

[0008] As a further improvement of the above technical solution, the radial distance between adjacent impact air intake structures is S1, and the radial dimension of the impact air intake structure is D1. S1 / D1 is in the range of 2.0 - 6.0.

[0009] As a further improvement of the above technical solution, the diameter of the flow disturbing structure on the suction side is D2, and the diameter of the flow disturbing structure on the pressure side is D3. The diameter of the flow disturbing structure is 0.3 - 0.6 mm; the distance L between two adjacent flow disturbing structures along the span of the blade is > D2 + D3.

[0010] As a further improvement of the above technical solution, the inlet height of the second chamber is F1, the outlet height is F2, and the height of the second chamber gradually decreases from the inlet end to the outlet end; the height of the flow disturbing structure is H, and the height of the flow disturbing structure gradually decreases from the inlet end to the outlet end to adapt to the height change in the second chamber.

[0011] As a further improvement of the above technical solution, the heights of the flow disturbing structure from the inlet end to the outlet end of the second chamber are H1 - Hn in sequence. F1 / H1 is in the range of 0.3 - 0.7, and F2 / Hn is in the range of 0.3 - 0.7.

[0012] As a further improvement of the above technical solution, the boss structure is in a cylindrical shape, a hemispherical shape or a frustum shape.

[0013] According to another aspect of the present invention, there is also provided a design method, including: S1. Determine the gas-side heat transfer boundary in the blade trailing edge region according to the overall parameters, and select a matching blade material; S2. Determine the metal limiting temperature and temperature gradient limiting value in the blade trailing edge region according to the selected blade material; S3. Preliminary planning and design of the flow disturbance structure and impingement air bleeding structure in the blade trailing edge region; S4. Obtain the heat transfer boundary on the cold air side of the trailing edge region through simulation calculation to obtain the temperature distribution of the blade trailing edge. If the highest temperature value is greater than the limiting value or the temperature gradient is greater than the limiting value, optimize the local dimensions and repeat step S4.

[0014] As a further improvement of the above technical solution, step S4 further includes: if the highest temperature value is less than the limiting value and the temperature gradient is less than the limiting value, go to step S5; Step S5 includes: adjusting the local dimensions within a preset range according to the process requirements.

[0015] According to another aspect of the present invention, an aeroengine is further provided, which includes the above turbine blade trailing edge cooling structure.

[0016] The present invention has the following beneficial effects: The turbine blade trailing edge cooling structure of the present invention is provided with a first chamber connected to the blade inner cavity and a second chamber connected to the blade trailing edge exhaust end and the first chamber at the trailing edge of the blade. After the cooling air in the blade inner cavity enters the first chamber from the inner cavity, it enters the second chamber through the impact air entrainment structure between the first chamber and the second chamber, flows through the spoiler structure on the blade basin side and the spoiler structure on the back side of the blade, and flows out into the main combustion gas through the slit of the trailing edge. Compared with the spoiler column array structure in the prior art, the impact air entrainment and boss combined cooling structure of the present structure can effectively enhance the local enhanced heat exchange. Introducing cooling air through the impact air entrainment structure for cooling can make high-speed cold air directly impact the surface of the blade inner cavity, generate high-intensity convective heat exchange at the part that needs to be cooled, and thus effectively reduce the local high temperature; and the boss structure can cause disturbance in the boundary layer, so that the heat exchange between the cooling air and the wall is more sufficient, thereby further enhancing the local heat exchange effect. Compared with the existing spoiler column, the strengthening effect of the present structure is more concentrated, and according to the heat load of different parts of the blade, the impact air entrainment structure and / or the boss structure can be adjusted. The distribution and size parameters of the boss structure can realize the reasonable distribution of cooling air, thereby further improving the cooling efficiency and making the flow distribution uniform and reasonable; by setting the impact air ducting structure, the cooling air can also enter the second chamber more orderly, and the boss structure has relatively little interference with the airflow, and the overall flow resistance is relatively low, avoiding the high length and strong resistance of flow turbulence caused by the formation of complex vortices and wakes due to the use of spoiler columns; and the boss structures staggered on the blade basin side and the blade back side can guide the airflow to a certain extent, so that the airflow is more in line with the inner cavity wall of the blade and thus reduce the airflow separation phenomenon, avoiding the large-scale separation behind the column caused by the use of the spoiler column structure, affecting the cooling effect and stability; the cooling structure is easier to adapt to the complex inner cavity shape of the blade. Compared with the spoiler column, the boss structure is relatively simple, the manufacturing difficulty and cost are reduced, and it can be realized through a variety of processes such as casting and machining, and the precision requirements are relatively low, especially in the narrow area of the trailing edge of the blade, the boss structure can be more flexibly arranged, which greatly reduces the processing difficulty and cost, and improves the processing efficiency.

[0017] In addition to the above-described purposes, features and advantages, the present invention has other purposes, features and advantages. The present invention will be further described in detail with reference to the accompanying drawings. BRIEF DESCRIPTION OF THE DRAWINGS

[0018] The drawings constituting a part of this application are used to provide a further understanding of the present invention. The exemplary embodiments of the present invention and their descriptions are used to explain the present invention and do not constitute an improper limitation of the present invention. In the drawings: Figure 1 is a schematic diagram of the blade structure of a preferred embodiment of the present invention; Figure 2 is a schematic structural diagram of a cooling structure of a preferred embodiment of the present invention; Figure 3 is a partial sectional view of the cooling structure of a preferred embodiment of the present invention; Figure 4 is a schematic diagram showing the distribution of the boss structures in the first embodiment of the present invention.

[0019] Legend: 100, the first chamber; 200, the second chamber; 300, the impact air bleeding structure; 400, the flow disturbance structure; 500, the trailing edge slit. Detailed implementation manners

[0020] The embodiments of the present invention will be described in detail below with reference to the accompanying drawings. However, the present invention can be implemented in many different ways defined and covered by the following.

[0021] Figure 1 is a schematic diagram of the blade structure of a preferred embodiment of the present invention; Figure 2 is a schematic diagram of the structure of the cooling structure of a preferred embodiment of the present invention; Figure 3 is a partial sectional view of the cooling structure of a preferred embodiment of the present invention.

[0022] As Figures 1 to 3 shown, in the trailing edge cooling structure of the turbine blade of this embodiment, the trailing edge of the blade has a first chamber 100 communicating with the inner cavity of the blade, and a second chamber 200 communicating with the exhaust end of the trailing edge of the blade and the first chamber 100 respectively. An impact air bleeding structure 300 is provided between the first chamber 100 and the second chamber 200. A flow disturbance structure 400 is provided on the inner wall of the second chamber 200. The flow disturbance structure 400 is a boss structure formed on the inner wall of the second chamber 200. The radial layout position of the flow disturbance structure 400 matches the radial layout position of the impact air bleeding structure 300. The flow disturbance structures 400 on the inner wall of the blade basin side and the inner wall of the blade back side of the second chamber 200 are arranged alternately.

[0023] It can be understood that the trailing edge cooling structure of the turbine blade is provided with a first chamber 100 connected to the inner cavity of the blade and a second chamber 200 connected to the exhaust end of the trailing edge of the blade and the first chamber 100 at the trailing edge of the blade. After the cooling air in the inner cavity of the blade enters the first chamber 100 from the inner cavity, it enters the second chamber 200 through the impact air bleed structure 300 between the first chamber 100 and the second chamber 200, flows through the spoiler structure 400 on the blade basin side and the spoiler structure 400 on the back side of the blade, and flows out into the main combustion gas through the trailing edge slit 500. The impact air bleed and boss combined cooling of the structure Compared with the spoiler column array structure in the prior art, the structural form can effectively enhance the local enhanced heat exchange. The cooling air introduced through the impact air bleed structure 300 can make the high-speed cold air directly impact the inner cavity surface of the blade, and generate high-intensity convective heat exchange at the position that needs to be cooled, thereby effectively reducing the local high temperature; and the boss structure can cause disturbances in the boundary layer, so that the heat exchange between the cooling air and the wall surface is more sufficient, thereby further enhancing the local heat exchange effect. Compared with the existing spoiler column, the strengthening effect of this structural form is more concentrated, and according to the heat load of different parts of the blade, The cooling air can also be reasonably distributed by adjusting the distribution and size parameters of the impact air induction structure 300 and / or the boss structure, thereby further improving the cooling efficiency and making the flow distribution uniform and reasonable; the impact air induction structure 300 can also make the cooling air enter the second chamber 200 more orderly, and the boss structure has relatively little interference with the airflow, and the overall flow resistance is relatively low, avoiding the formation of complex vortices and wakes due to the use of spoiler columns. The flow turbulence has a long length and strong resistance; and the boss structures staggered on the blade basin side and the blade back side can induce The airflow is guided so that the airflow is closer to the inner cavity wall of the blade, thereby reducing the airflow separation phenomenon, avoiding the large-scale separation behind the column caused by the use of a spoiler column structure, which affects the cooling effect and stability; the cooling structure is easier to adapt to the complex inner cavity shape of the blade. Compared with the spoiler column, the boss structure is relatively simple, and the manufacturing difficulty and cost are reduced. It can be realized through a variety of processes such as casting and machining, and the precision requirements are relatively low. Especially in the narrow area of the trailing edge of the blade, the boss structure can be more flexibly arranged, which greatly reduces the processing difficulty and cost and improves the processing efficiency.

[0024] It should be understood that the boss structure is preferably cylindrical, hemispherical or truncated cone-shaped, and the boss structures are consistent in shape, so that the structure is regular and simple, and the processing difficulty is low.

[0025] In some embodiments, the impact air bleed structure 300 includes a plurality of impact holes uniformly distributed along the radial direction of the blade, wherein the radial distribution position of each boss structure matches the impact air bleed structure 300, that is, the first row of spoiler structures 400 is located at the rear end of the impact air bleed structure, so that the airflow induced by the impact air bleed structure 300 impacts the boss surface to enhance heat exchange; Further, the radial spacing between adjacent impact air bleeding structures 300 is S1, and the radial dimension of the impact air bleeding structure 300 is D1. S1 / D1 is in the range of 2.0 - 6.0. Within the above range, the cooling effects of the impact holes and the boss structures can be ensured under the condition of limited cold air volume. It should be noted that the diameter D1 of the impact holes is matched and designed according to the cold air flow distribution requirements of the blade and the wall thickness of the rear cavity of the blade. For example, in a specific embodiment, D1 = 0.45 mm and S1 = 2.25 mm.

[0026] In some embodiments, the diameter of the spoiler structure 400 on the suction side of the blade is D2, and the diameter of the spoiler structure 400 on the pressure side of the blade is D3. Among them, the diameter of the spoiler structure 400 on the suction side of the blade and the diameter of the spoiler structure 400 on the pressure side of the blade are selected within the range of 0.3 - 0.6 mm. It can be understood that due to the limitations of the existing casting process and structural strength, the diameter of the boss structure is controlled above 0.3 mm to ensure the feasibility of casting processing while meeting the requirements of structural strength design. Moreover, in order to arrange as many boss structures as possible in the limited space of the second chamber 200 to enhance heat transfer, the diameter of the boss structure should not be too large. Therefore, based on the design dimensions of the aeroengine turbine blade and relevant engineering experience, the boss diameter is controlled within 0.3 mm - 0.6 mm. Furthermore, on the premise of meeting the requirements of strength design and casting process, a strong heat transfer effect can be ensured in the limited chamber space. It should be noted that the spacing L between two adjacent spoiler structures 400 along the blade span is greater than D2 + D3. In the span direction, multiple rows of boss structures are arranged on the pressure side and the suction side of the blade respectively, and they are arranged in a cross pattern. The number of rows of the boss structures increases or decreases according to the area size and cooling requirements of the second chamber 200 in the trailing edge region of the blade. The spanwise spacing between the center points of the adjacent boss structures on the pressure side and the suction side of the blade is greater than D2 + D3 to prevent the cold air from being insufficiently disturbed between the bosses on the suction side and the pressure side of the blade and reducing the cooling effect.

[0027] In some embodiments, the inlet height of the second chamber 200 is F1, the outlet height is F2, and the height of the second chamber 200 gradually decreases from the inlet end to the outlet end; the height of the spoiler structure 400 is H, and the height of the spoiler structure 400 gradually decreases in the direction from the inlet end to the outlet end to adapt to the height change in the second chamber 200. It should be understood that generally, the inner cavity of the blade will become narrower and narrower towards the trailing edge of the blade. Therefore, F2 < F1. In order to adapt to the change in the cross-sectional area of the inner cavity at the trailing edge of the blade, the height of the spoiler structure 400 changes adaptively with the chamber height; the heights of the spoiler structure 400 from the inlet end to the outlet end of the second chamber 200 are H1-Hn in sequence, that is, H1 > H2 >... > Hn; among them, F1 and F2 are jointly determined by the blade profile wall thickness in the trailing edge area of the blade and the wall thickness requirements of the blade basin and blade back in the trailing edge area. According to the engineering design experience of engine turbine blades, F1 and F2 are greater than 0.5 mm; F1 / H1 is within the range of 0.3 to 0.7, and F2 / Hn is within the range of 0.3 to 0.7. Among them, F1 / H1 is greater than 0.3 to ensure that the boss structure has a strong heat transfer effect, but the height of the boss structure should not be too large, and it is controlled within 0.7 to prevent a sudden increase in the cold air flow resistance. The range control of F2 / Hn is the same. When designing within this range, the flow resistance of the boss structure is small and the heat transfer effect is strong.

[0028] Based on the above technical solution, the shape of the spoiler structure 400 is preferably cylindrical, that is, the radial dimensions of each boss structure are the same, the radial dimensions and height dimensions of each row of boss structures are the same, and there is only one variable in the height dimension, so as to further reduce the processing difficulty and improve the processing efficiency.

[0029] On the other hand, a preferred embodiment of the present invention also provides a design method, including: S1. Determine the heat transfer boundary on the gas side in the trailing edge area of the blade according to the overall parameters, and select a matching blade material; S2. Determine the metal limiting temperature and temperature gradient limiting value in the trailing edge area of the blade according to the selected blade material; S3. Initially plan and design the spoiler structure 400 and the impingement air supply structure 300 in the trailing edge area of the blade; S4. Obtain the heat transfer boundary on the cold air side in the trailing edge area through simulation calculation to obtain the temperature distribution at the trailing edge of the blade. If the highest temperature value is greater than the limiting value or the temperature gradient is greater than the limiting value, optimize the local dimensions and repeat step S4; if the highest temperature value is less than the limiting value and the temperature gradient is less than the limiting value, go to step S5; S5. Adjust the local dimensions within the preset range according to the process requirements.

[0030] This design method selects the blade material according to the matching of the heat transfer boundary on the gas side of the blade trailing edge region, determines the metal limiting temperature and temperature gradient limiting value in the blade trailing edge region based on the selected blade material, and then can preliminarily plan and design the turbulator structure 400 and the impingement air bleeding structure 300 in the blade trailing edge region of this embodiment according to design manuals, engineering experience, etc. Based on the preliminary design scheme, simulation calculations are carried out, and the local dimensions are adjusted and optimized according to the limiting temperature and temperature gradient limit obtained in step S2 for iterative design, so that the highest temperature and temperature gradient are both less than the corresponding limiting values. The design method is concise and reasonable. The finally obtained combined cooling structure can significantly strengthen local heat transfer and effectively improve heat transfer uniformity, with improved overall structural strength and more reasonable structure.

[0031] It should be understood that in step S5, the local dimensions are adjusted within a reasonable preset range according to the processing technology requirements, so as to change the processing while avoiding affecting the heat transfer effect, structural strength, etc.

[0032] On the other hand, a preferred embodiment of the present invention also provides an aeroengine, which is applied with the above turbine blade trailing edge cooling structure.

[0033] Embodiment 1 In this embodiment, with reference to Figure 4 , the X direction is the spanwise direction of the blade, the Y direction is the radial direction. There are a total of 6 rows of boss structures arranged in the spanwise direction. Among them, 3 rows of boss structures are distributed on the suction side of the blade, and 3 rows of boss structures are distributed on the pressure side of the blade. In other embodiments, the number of rows of boss structures can be appropriately increased or decreased according to the area size of the second chamber 200 in the trailing edge region and the cooling requirements; the first row of boss structures is located on the suction side of the blade, and the distance between its centerline in the spanwise direction and the impingement holes is L1. The second row of boss structures is located on the pressure side of the blade, and the distance from the first row of boss structures is L2. The distances between the subsequent adjacent rows of boss structures are L3, L4, L5, and L6 in sequence; and preferably, D2 = D3, L1 = L2 = L3 = L4 = L5 = L6 > (D2 + D3). When the air flow passes through the bosses, it can be fully turbulized to strengthen heat transfer, and at the same time, the parameter consistency is high, which is more convenient for calculation and processing; specifically, in this embodiment, D2 = D3 = 0.3 mm, L = 0.8 mm, which are evenly arranged in the limited space of the rear chamber of the blade, and ensure a strong cooling effect; In the radial direction, in order to enable the air flow flowing out of the impact holes to directly impact the surface of the boss structure to enhance heat transfer, the boss structure is arranged directly behind the impact holes. The radial distance between the impact holes and the first row of boss structures is S1, and the diameter of the impact holes is D1. To ensure the cooling effects of the impact and the bosses under the condition of limited cold air volume, S1 / D1 is controlled within the range of 2.0 - 6.0. Among them, the diameter D1 of the impact holes is determined according to the air flow distribution requirements of the blade cold air and the wall thickness of the rear cavity of the blade. In this embodiment, D1 = 0.45 mm and S1 = 2.25 mm. The height of the inlet end of the second chamber 200 is F1, the height of the outlet end is F2, the height of the first row of bosses on the blade suction side and the pressure side is H1, the height of the second row of bosses on the blade suction side and the pressure side is H2, and the height of the third row of bosses on the blade suction side and the pressure side is H3. Towards the trailing edge of the blade, the chambers in the inner cavity of the blade become narrower, that is, F2 < F1, and the boss height also changes adaptively according to the cross-sectional area of the second chamber 200, that is, H1 > H2 > H3. And, F1 / H1 is within the range of 0.3 - 0.7, and F2 / H3 is within the range of 0.3 - 0.7, ensuring the high cooling capacity of the combined cooling structure and taking into account the flow resistance of the inner cavity in the trailing edge area. Specifically, F1 = 0.6 mm, F2 = 0.5 mm, H1 = 0.2 mm, H2 = 0.18 mm, H3 = 0.15 mm. When applied to this embodiment, the flow resistance of the bosses is small and the heat transfer effect is strong.

[0034] In this embodiment, the impinging bleeding directly injects the cooling air flow onto the inner cavity surface of the blade. Near the impinging point, the heat transfer coefficient can be increased by several times or even higher compared with ordinary convective cooling. In experimental studies, the local heat transfer coefficient of impingement cooling can reach 2000 - 5000 W / (m2•K), while the local heat transfer coefficient of simple pin fin cooling is generally in the range of 1000 - 3000 W / (m2•K). Moreover, the boss structure can generate additional disturbances within the boundary layer, further enhancing the heat transfer effect, reducing the thermal resistance within the boundary layer. Under the combined action, the overall heat transfer effect can be improved by 3% - 15%. Due to the arrangement and flow characteristics of the pin fins, the cold air flow rate in the channel is unevenly distributed radially, and the wall heat transfer shows the characteristics of being low in the middle and high on both sides. The combined structure of this embodiment makes the cooling air flow cover the inner cavity surface of the blade more evenly by reasonably designing the layout of the impinging holes and the shape and position of the bosses, reducing the non-uniformity of heat transfer and increasing the heat transfer uniformity by 10% - 20%. In addition, since the pin fins are columnar structures, when the blade is subjected to loads such as high temperature and high centrifugal force, stress concentration is likely to occur at the connection part between the pin fins and the blade wall. The boss structure of this embodiment is relatively smooth in connection with the blade, which is more conducive to stress transmission, reducing the stress concentration phenomenon, and can reduce the local stress concentration coefficient by 5% - 10%. Without adding too much weight to the blade, this structure provides a certain support and strengthening effect for the inner cavity of the blade. When subjected to the same centrifugal force and thermal load as the pin fin structure, the blade has a smaller deformation amount, improving the overall structural stability and reliability of the blade.

[0035] In the description of the present invention, it should be noted that the orientation or positional relationship indicated by the terms "upper", "lower", "front", "rear", etc. is based on the orientation or positional relationship shown in the drawings. It is only for the convenience of describing the present invention and simplifying the description, rather than indicating or implying that the device or element referred to must have a specific orientation, be constructed and operated in a specific orientation, and therefore should not be construed as a limitation to the present invention.

[0036] In the description of the present invention, it should be noted that unless otherwise clearly specified and defined, the terms "installed", "connected", and "connected" should be understood in a broad sense. For example, it can be a fixed connection, a detachable connection, or an integral connection; it can be a mechanical connection or an electrical connection; it can be directly connected or indirectly connected through an intermediate medium, and it can be the communication inside two elements. For those of ordinary skill in the art, the specific meanings of the above terms in the present invention can be understood according to specific situations.

[0037] The above are only the preferred embodiments of the present invention and are not intended to limit the present invention. For those skilled in the art, the present invention can have various modifications and changes. Any modification, equivalent replacement, improvement, etc. made within the spirit and principle of the present invention shall be included within the protection scope of the present invention.

Claims

1. A trailing edge cooling structure of a turbine blade, characterized in that, The trailing edge of the blade has a first chamber (100) communicating with the inner cavity of the blade, and a second chamber (200) communicating with the exhaust end of the trailing edge of the blade and the first chamber (100) respectively. An impact air bleeding structure (300) is arranged between the first chamber (100) and the second chamber (200). The inner wall of the second chamber (200) is provided with a flow disturbance structure (400). The flow disturbance structure (400) is a boss structure formed on the inner wall of the second chamber (200). The radial layout position of the flow disturbance structure (400) matches the radial layout position of the impact air bleeding structure (300). The flow disturbance structures (400) on the inner wall of the blade basin side and the inner wall of the blade back side of the second chamber (200) are arranged in a staggered manner.

2. The trailing edge cooling structure of the turbine blade according to claim 1, characterized in that, The impact air bleeding structure (300) includes a plurality of impact holes uniformly distributed along the radial direction of the blade.

3. The trailing edge cooling structure of the turbine blade according to claim 2, wherein The radial spacing between adjacent impact air bleeding structures (300) is S1, the radial dimension of the impact air bleeding structure (300) is D1, and S1 / D1 is in the range of 2.0 - 6.

0.

4. The trailing edge cooling structure of the turbine blade according to claim 1, wherein, The diameter of the flow disturbance structure (400) on the blade basin side is D2, the diameter of the flow disturbance structure (400) on the blade back side is D3, and the diameter of the flow disturbance structure (400) is 0.3 - 0.6 mm; the spacing L between two adjacent flow disturbance structures (400) along the blade span is > D2 + D3.

5. The trailing edge cooling structure of the turbine blade according to claim 1, characterized in that, The inlet height of the second chamber (200) is F1, the outlet height is F2, and the height of the second chamber (200) gradually decreases from the inlet end to the outlet end; the height of the flow disturbance structure (400) is H, and the height of the flow disturbance structure (400) gradually decreases from the inlet end to the outlet end direction to adapt to the height change in the second chamber (200).

6. The trailing edge cooling structure of the turbine blade according to claim 5, characterized in that, The heights of the flow disturbance structure (400) from the inlet end to the outlet end direction of the second chamber (200) are H1 - Hn in sequence, and F1 / H1 is in the range of 0.3 - 0.7, and F2 / Hn is in the range of 0.3 - 0.

7.

7. The trailing edge cooling structure of the turbine blade according to claim 1, characterized in that, The boss structure is cylindrical or hemispherical or frustum-shaped.

8. A design method, characterized in that, Including: S1. Determine the gas-side heat transfer boundary in the trailing edge area of the blade according to the overall parameters, and select a matching blade material; S2. Determine the metal limit temperature and temperature gradient limit value in the trailing edge area of the blade according to the selected blade material; S3. Preliminary planning and design of the flow disturbance structure (400) and impact air bleeding structure (300) in the trailing edge area of the blade; S4. Obtain the heat transfer boundary on the cold air side in the trailing edge area through simulation calculation to obtain the temperature distribution of the blade trailing edge. If the highest temperature value is greater than the limit value or the temperature gradient is greater than the limit value, optimize the local dimensions and repeat step S4.

9. The design method according to claim 8, wherein Step S4 further includes: if the highest temperature value is less than the limit value and the temperature gradient is less than the limit value, go to step S5; Step S5 includes: adjusting the local dimensions within a preset range according to the process requirements.

10. An aeroengine, characterized in that, The turbine blade trailing edge cooling structure according to any one of claims 1 - 7 is applied.

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