Turbine blade trailing edge cooling structure, design method and aircraft engine

Abstract

The present invention discloses a turbine blade trailing edge cooling structure, design method, and aircraft engine. The trailing edge of the blade has a first chamber connected to the blade inner cavity and a second chamber connected to the exhaust end of the blade trailing edge and the first chamber, respectively. An impact bleed structure is provided between the first and second chambers. The inner wall of the second chamber is provided with a turbulent structure. The turbulent structure is a boss structure formed on the inner wall of the second chamber. The radial layout position of the turbulent structure matches the radial layout position of the impact bleed structure. The turbulent structure on the inner wall of the blade basin side of the second chamber and the turbulent structure on the inner wall of the blade back side are staggered. Compared with the turbulent column array structure, this structure effectively enhances heat exchange. The impact bleed causes high-speed cold air to impact the inner cavity surface of the blade, generating high-intensity convective heat exchange in the area requiring focused cooling, effectively reducing local high temperature. The boss can cause disturbances in the boundary layer, making the heat exchange between the cold air and the wall more complete, further enhancing the local heat exchange effect.

Description

Technical Field

[0001] The present invention relates to the technical field of turbine blades, and in particular to a turbine blade trailing edge cooling structure. In addition, the present invention also relates to an aircraft engine comprising the turbine blade trailing edge cooling structure. Background Art

[0002] With the continuous advancement of high-performance aircraft engine and gas turbine technology, engine thermodynamic cycle parameters continue to push the boundaries. This trend has made thermal protection of turbine blades, a critical hot-end component, increasingly challenging. Blades are constantly exposed to the harsh environment of high temperature, high pressure, and high-velocity airflow, requiring stringent thermal protection measures. Air cooling technology is the primary means of addressing this thermal protection challenge for turbine blades. By meticulously planning the cooling air flow path within the blades, a highly complex cooling airflow system is constructed to ensure stable operation even under high-temperature conditions.

[0003] The basic cooling methods for aircraft engine turbine blades are mainly convection cooling, impingement cooling, and film cooling. In actual 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 at the trailing edge of the turbine blade, an array of spoiler columns is often provided. These spoiler columns connect the back of the blade and the blade basin in one or more rows, and their effect is significant. On the one hand, the spoiler columns can cause the cold air to form strong convection in the spoiler column area, greatly improving the heat exchange capacity. On the other hand, it 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 lot of practice and optimization, the spoiler column array enhanced cooling structure has become very mature. This structure has been widely used in many engines currently under development and in service.

[0004] The main shortcomings of the above technical solution are as follows:

[0005] 1) Under the conditions of limited space and limited cooling air volume, the cooling effect of the spoiler array cooling structure has a clear upper limit, and it is difficult to improve the cooling effect of the trailing edge area by simply changing the spoiler parameters;

[0006] 2) The cold air flow resistance in the area of ​​a single spoiler array is very large, causing the airflow to form complex vortices and wakes around the spoilers, increasing the turbulence and resistance of the flow. This also requires a high cold air supply pressure to ensure sufficient cold air flow, which is very costly to achieve in the engine.

[0007] 3) The shape, structure and arrangement are complex, easily restricted by the narrow space at the trailing edge of the blade, and have high requirements on manufacturing process and mold precision. Summary of the Invention

[0008] The present invention provides a turbine blade trailing edge cooling structure, a design method and an aircraft engine to solve the technical problems in the prior art of the spoiler column array structure, such as the difficulty in improving the cooling effect, the large flow resistance resulting in a large demand for cooling air and poor stability, and the complex structure resulting in poor adaptability and high processing difficulty.

[0009] According to one aspect of the present invention, a turbine blade trailing edge cooling structure is provided, wherein the trailing edge of the blade has a first chamber connected to the inner cavity of the blade and a second chamber connected to the exhaust end of the trailing edge of the blade and the first chamber respectively, an impact induced air structure is arranged between the first chamber and the second chamber, and an inner wall of the second chamber is provided with a turbulence structure, the turbulence structure is a boss structure formed on the inner wall of the second chamber, the radial arrangement position of the turbulence structure matches the radial arrangement position of the impact induced air structure, and the turbulence structure on the inner wall of the blade basin side of the second chamber and the turbulence structure on the inner wall of the blade back side are arranged alternately.

[0010] As a further improvement of the above technical solution, the impact air entrainment structure includes a plurality of impact holes evenly distributed along the radial direction of the blade.

[0011] As a further improvement of the above technical solution, the radial spacing between adjacent impact air entrainment structures is S1, the radial dimension of the impact air entrainment structure is D1, and S1 / D1 is in the range of 2.0-6.0.

[0012] As a further improvement of the above technical solution, the diameter of the spoiler structure on the blade basin side is D2, and the diameter of the spoiler structure on the blade back side is D3. The diameter of the spoiler structure is 0.3-0.6mm; the distance L between two adjacent spoiler structures along the blade span is greater than D2+D3.

[0013] 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 spoiler structure is H, and the height of the spoiler structure gradually decreases from the inlet end to the outlet end to adapt to the height change in the second chamber.

[0014] As a further improvement of the above technical solution, the height of the spoiler structure from the inlet end to the outlet end of the second chamber is H1-Hn, H1 / F1 is in the range of 0.3 to 0.7, and Hn / F2 is in the range of 0.3 to 0.7.

[0015] As a further improvement of the above technical solution, the boss structure is cylindrical, hemispherical or truncated cone-shaped.

[0016] According to another aspect of the present invention, there is also provided a design method, comprising:

[0017] S1. Determine the gas-side heat exchange boundary of the blade trailing edge area based on overall parameters and select a matching blade material;

[0018] S2. Determine the metal limit temperature and temperature gradient limit value of the blade trailing edge area based on the selected blade material;

[0019] S3. Preliminary planning and design of the turbulence structure and impingement air bleed structure at the trailing edge of the blade;

[0020] S4. The heat exchange boundary on the cold air side of the trailing edge region is obtained through simulation calculation to obtain the temperature distribution of the blade trailing edge. If the maximum temperature value is greater than the limit value or the temperature gradient is greater than the limit value, repeat step S4 for local size optimization.

[0021] As a further improvement of the above technical solution, step S4 further includes: if the maximum temperature value is less than the limit value and the temperature gradient is less than the limit value, go to step S5;

[0022] Step S5 includes: adjusting the local size within a preset range according to process requirements.

[0023] According to another aspect of the present invention, an aircraft engine is provided, which includes the above-mentioned turbine blade trailing edge cooling structure.

[0024] The present invention has the following beneficial effects:

[0025] 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 exhaust end of the blade trailing edge 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 bleed structure between the first and second chambers, flows through the turbulent flow structure on the blade basin side and the turbulent flow structure on the blade back side, 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 bleed and boss combined cooling structure of this structure can effectively enhance local enhanced heat exchange. Introducing cooling air through the impact bleed structure for cooling can cause high-speed cold air to directly impact the blade inner cavity surface, generating high-intensity convective heat exchange in the area requiring focused cooling, thereby effectively reducing the local high temperature. The boss structure can cause disturbances in the boundary layer, making the heat exchange between the cooling air and the wall more sufficient, thereby further enhancing the local heat exchange effect. Compared with the existing spoiler column structure, the strengthening effect of this structure is more concentrated. In addition, according to the heat load of different parts of the blade, the impact bleed structure and / or boss can be adjusted. The distribution and size parameters of the platform structure 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 induction 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 application 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; this 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 blade trailing edge, the boss structure can be more flexibly arranged, greatly reducing the processing difficulty and cost, and improving the processing efficiency.

[0026] In addition to the above-described objects, features and advantages, the present invention has other objects, features and advantages. The present invention will be further described in detail below with reference to the accompanying drawings. BRIEF DESCRIPTION OF THE DRAWINGS

[0027] The accompanying drawings, which constitute part of this application, are intended to provide a further understanding of the present invention. The exemplary embodiments of the present invention and their descriptions are intended to explain the present invention and do not constitute an undue limitation of the present invention. In the accompanying drawings:

[0028] Figure 1 Schematic diagram of the blade structure of a preferred embodiment of the present invention;

[0029] Figure 2is a schematic structural diagram of a cooling structure according to a preferred embodiment of the present invention;

[0030] Figure 3 is a partial cross-sectional view of a cooling structure according to a preferred embodiment of the present invention;

[0031] Figure 4 Schematic diagram of the boss structure distribution of embodiment 1 of the present invention.

[0032] Legend:

[0033] 100. First chamber; 200. Second chamber; 300. Impact air bleed structure; 400. Turbine structure; 500. Trailing edge slit. DETAILED DESCRIPTION

[0034] The embodiments of the present invention are described in detail below with reference to the accompanying drawings. However, the present invention can be implemented in many different ways as defined and covered below.

[0035] Figure 1 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 according to a preferred embodiment of the present invention; Figure 3 It is a partial cross-sectional view of the cooling structure of a preferred embodiment of the present invention.

[0036] like Figures 1 to 3 As shown, the turbine blade trailing edge cooling structure of this embodiment has a first chamber 100 connected to the inner cavity of the blade and a second chamber 200 respectively connected to the exhaust end of the trailing edge of the blade and the first chamber 100. An impact induced air 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 spoiler structure 400. The spoiler structure 400 is a boss structure formed on the inner wall of the second chamber 200. The radial arrangement position of the spoiler structure 400 matches the radial arrangement position of the impact induced air structure 300. The spoiler structure 400 on the inner wall of the blade basin side of the second chamber 200 and the spoiler structure 400 on the inner wall of the blade back side are arranged alternately.

[0037] It can be understood that the turbine blade trailing edge cooling structure is provided with a first chamber 100 connected to the blade inner cavity and a second chamber 200 connected to the blade trailing edge exhaust end and the first chamber 100 at the trailing edge of the blade. After the cooling air in the blade inner cavity enters the first chamber 100 from the inner cavity, it enters the second chamber 200 through the impact bleed structure 300 between the first chamber 100 and the second chamber 200, flows through the turbulence structure 400 on the blade basin side and the turbulence structure 400 on the blade back side, and flows out into the main combustion gas through the trailing edge slit 500. The impact bleed and boss combined cooling of this 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 directly impact the inner cavity surface of the blade, generating high-intensity convective heat exchange at the parts that need to be cooled, thereby effectively reducing the local high temperature; and the boss structure can cause disturbances in the boundary layer, making the heat exchange between the cooling air and the wall 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 reasonable distribution of cooling air can also be achieved 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; by setting the impact air induction structure 300, the cooling air can also enter the second chamber 200 in a more orderly manner, and the boss structure has relatively little interference with the air flow, and the overall flow resistance is relatively low, avoiding the flow turbulence caused by the application of the spoiler column to form complex vortices and the tail flow having a high length and strong resistance; and the boss structures staggered on the blade basin side and the blade back side can to a certain extent induce The airflow is guided so that the airflow fits more closely 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 the 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 various 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 arranged more flexibly, which greatly reduces the processing difficulty and cost and improves the processing efficiency.

[0038] It should be understood that the boss structure is preferably cylindrical, hemispherical or truncated cone-shaped, and the shapes of the boss structures are consistent, so that the structure is regular and simple, and the processing difficulty is low.

[0039] In some embodiments, the impingement air bleed structure 300 includes a plurality of impingement holes uniformly distributed along the radial direction of the blade, wherein the radial distribution position of each boss structure matches the impingement air bleed structure 300, that is, the first row of flow-turbine structures 400 is located at the rear end of the impingement air bleed structure, so that the airflow induced by the impingement air bleed structure 300 impinges on the boss surface to enhance heat exchange;

[0040] Furthermore, the radial spacing between adjacent impact air bleed structures 300 is S1, the radial dimension of the impact air bleed structure 300 is D1, and S1 / D1 is in the range of 2.0-6.0. Within the above range, the cooling effect of the impact hole and the boss structure can be guaranteed under the condition of limited cooling air volume; it should be noted that the impact hole diameter D1 is designed to match the blade cooling air flow distribution requirements and the wall thickness of the blade rear cavity. For example, in a specific embodiment, D1=0.45mm, S1=2.25mm.

[0041] In some embodiments, the diameter of the spoiler structure 400 on the blade basin side is D2, and the diameter of the spoiler structure 400 on the blade back side is D3, wherein the diameter of the spoiler structure 400 on the blade basin side and the diameter of the spoiler structure 400 on the blade quilt side are selected in the range of 0.3-0.6 mm. It can be understood that, limited by the existing casting process and structural strength, the diameter of the boss structure is controlled to be above 0.3 mm, so as to meet the structural strength design requirements while ensuring the feasibility of the casting process, and in order to arrange as many boss structures as possible in the limited space of the second chamber 200 to enhance heat exchange, the diameter of the boss structure should not be too large. Therefore, based on the design size of the aircraft engine turbine blade and related engineering experience, According to the experiment, the boss diameter is controlled within the range of 0.3mm to 0.6mm, thereby ensuring a strong heat exchange effect in a limited chamber space while meeting the strength design requirements and casting process requirements; it should be noted that the spacing L between two adjacent spoiler structures 400 along the span direction of the blade is greater than D2+D3, and in the span direction, multiple rows of boss structures are respectively provided on the back side of the blade and the blade basin side, and are arranged crosswise, and the number of rows of boss structures increases or decreases according to the area of ​​the second chamber 200 in the trailing edge area of ​​the blade and the cooling demand; the span-wise spacing between the center points of the boss structures on the adjacent back side and blade basin side is greater than D2+D3 to prevent the cold air from being insufficiently turbulent between the blade basin and the back boss and reducing the cooling effect.

[0042] 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 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 becomes narrower towards the trailing edge of the blade, so F2 < F1. 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 height of the chamber; 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; wherein, F1 and F2 are jointly determined by the blade profile wall thickness in the trailing edge region of the blade and the requirements for the wall thickness of the blade basin and back in the trailing edge region. According to the engineering design experience of engine turbine blades, F1 and F2 are greater than 0.5 mm; H1 / F1 is within the range of 0.3 - 0.7, and Hn / F2 is within the range of 0.3 - 0.7. Among them, H1 / F1 is greater than 0.3 to ensure that the boss structure has a strong heat exchange 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 control range of Hn / F2 is the same. Designed within this range, the boss structure has a small flow resistance and a strong heat exchange effect.

[0043] 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.

[0044] On the other hand, a preferred embodiment of the present invention also provides a design method, including:

[0045] S1. Determine the gas-side heat transfer boundary in the trailing edge region of the blade according to the overall parameters, and select a matching blade material;

[0046] S2. Determine the metal limiting temperature and temperature gradient limiting value in the trailing edge region of the blade according to the selected blade material;

[0047] S3. Preliminary plan and design the spoiler structure 400 and the impact air bleeding structure 300 in the trailing edge region of the blade;

[0048] S4. Obtain the heat transfer boundary on the cold air side in the trailing edge region 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;

[0049] S5. Adjust the local dimensions within the preset range according to the process requirements.

[0050] This design method selects blade materials based on the gas-side heat exchange boundary matching of the blade trailing edge area, and determines the metal limit temperature and temperature gradient limit value of the blade trailing edge area based on the selected blade material. Then, the spoiler structure 400 and the impact air bleed structure 300 of the blade trailing edge area of ​​this embodiment can be preliminarily planned and designed based on the design manual, engineering experience, etc., and simulation calculations are performed based on the preliminary design scheme. According to the limit temperature and temperature gradient limit obtained in step S2, the local dimensions are adjusted and optimized for iterative design, so that the maximum temperature and temperature gradient are both less than the corresponding limit values. The design method is streamlined and reasonable, and the resulting combined cooling structure can significantly enhance local heat exchange and effectively improve heat exchange uniformity, and the overall structural strength is improved and the structure is more reasonable.

[0051] It should be understood that in step S5, the local dimensions are adjusted within a reasonable preset range according to the processing requirements, so as to make the processing more variable while avoiding affecting the heat exchange effect, structural strength, etc.

[0052] On the other hand, a preferred embodiment of the present invention further provides an aircraft engine, which is equipped with the above-mentioned turbine blade trailing edge cooling structure.

[0053] Example 1

[0054] In this embodiment, reference Figure 4 , the X direction is the span direction of the blade, the Y direction is the radial direction, and a total of 6 rows of boss structures are provided in the span direction, wherein 3 rows of boss structures are distributed on the blade basin side, and 3 rows of boss structures are distributed on the blade back side. In other embodiments, the number of rows of boss structures is appropriately increased or decreased according to the area of ​​the second chamber 200 in the trailing edge region and the cooling requirement; the first row of boss structures is located on the blade basin side, and the distance between it and the span centerline of the impact hole is L1, and the second row of boss structures is located on the blade back side, and the distance between it and the first row of boss structures is L2. The spacing between the boss structures in the adjacent rows is L2, and the spacing between the boss structures in the subsequent adjacent rows is L3, L4, L5, and L6, respectively. Furthermore, preferably, D2=D3, L1=L2=L3=L4=L5=L6>(D2+D3). When the airflow passes through the boss, it can be fully disturbed, thereby enhancing heat exchange. At the same time, the high degree of parameter consistency makes calculation and processing easier. Specifically, D2=D3=0.3mm and L=0.8mm in this embodiment are evenly arranged in the limited space of the rear cavity of the blade, thereby ensuring a strong cooling effect.

[0055] 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. In order to ensure the cooling effects of the impact and the bosses under the condition of limited cooling air volume, S1 / D1 is controlled within the range of 2.0 to 6.0. Among them, the diameter D1 of the impact holes is determined according to the air flow distribution requirements of the blade cooling 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, H1 / F1 is within the range of 0.3 to 0.7, and H3 / F2 is within the range of 0.3 to 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 region. Specifically, F1 = 0.6 mm, F2 = 0.5 mm, H1 = 0.2 mm, H2 = 0.18 mm, H3 = 0.15 mm. Applied to this embodiment, the flow resistance of the bosses is small and the heat transfer effect is strong.

[0056] In this embodiment, the impingement air injection directly sprays the cooling airflow onto the inner cavity surface of the blade, which can increase the heat transfer coefficient several times or even higher near the impact point compared with ordinary convection cooling; in experimental research, the local heat transfer coefficient of impingement cooling can reach 2000~5000W / (m2·K), while the local heat transfer coefficient of simple spoiler cooling is generally 1000~3000W / (m2·K); in addition, the boss structure can generate additional disturbances in the boundary layer, further enhancing the heat transfer effect and reducing the thermal resistance in the boundary layer. Under the combined effect, the overall heat transfer effect can be improved by 3%~15%; due to the arrangement and flow characteristics of the spoiler columns, the cold air flow in the channel will be unevenly distributed in the radial direction, and the wall heat transfer will be low in the middle and high on both sides. The combined structure of this embodiment makes the cooling airflow cover the inner cavity surface of the blade more evenly by rationally designing the layout of the impact holes and the shape and position of the boss, thereby reducing the unevenness of heat exchange and improving the heat exchange uniformity by 10% to 20%; in addition, since the spoiler column is a columnar structure, when the blade is subjected to loads such as high temperature and high centrifugal force, stress concentration is easily generated at the connection between the spoiler column and the blade wall; the boss structure of this embodiment avoids a relatively smooth connection with the blade, which makes it easier to transfer stress and reduce stress concentration, and can reduce the local stress concentration coefficient by 5% to 10%; this structure provides certain support and reinforcement for the inner cavity of the blade without increasing too much weight of the blade. Compared with the spoiler column structure, when subjected to the same centrifugal force and thermal load, the blade deformation is smaller, thereby improving the overall structural stability and reliability of the blade.

[0057] In the description of the present invention, it should be noted that the terms "upper", "lower", "front", "back", etc. indicate orientations or positional relationships based on the orientations or positional relationships shown in the accompanying drawings. They are only for the convenience of describing the present invention and simplifying the description, and do not indicate or imply that the device or element referred to must have a specific orientation, be constructed and operated in a specific orientation. Therefore, they cannot be understood as limiting the present invention.

[0058] In the description of the present invention, it should be noted that, unless otherwise expressly specified or limited, the terms "mounted," "connected," and "connected" should be understood in a broad sense. For example, they may refer to fixed, detachable, or integral connections; mechanical or electrical connections; direct or indirect connections through an intermediate medium; and internal communication between two components. Those skilled in the art will understand the specific meanings of the above terms in the present invention based on the specific circumstances.

[0059] The foregoing description is merely a preferred embodiment of the present invention and is not intended to limit the present invention. Those skilled in the art will readily appreciate that various modifications and variations of the present invention are possible. Any modifications, equivalent substitutions, or improvements made within the spirit and principles of the present invention are intended to be within the scope of protection of the present invention.

Claims

1. A turbine blade trailing edge cooling structure, characterized in that: The trailing edge of the blade has a first chamber (100) connected to the inner cavity of the blade and a second chamber (200) respectively connected to the exhaust end of the trailing edge of the blade and the first chamber (100); an impact induced air structure (300) is provided between the first chamber (100) and the second chamber (200); an inner wall of the second chamber (200) is provided with a turbulence structure (400); the turbulence structure (400) is a boss structure formed on the inner wall of the second chamber (200); the radial arrangement position of the turbulence structure (400) matches the radial arrangement position of the turbulence structure (300); the turbulence structure (400) on the inner wall of the blade basin side of the second chamber (200) and the turbulence structure (400) on the inner wall of the blade back side are arranged in an alternating manner; the turbulence structure (400) on the blade basin side The diameter of the spoiler structure (400) is D2, the diameter of the spoiler structure (400) on the back side of the blade is D3, and the diameter of the spoiler structure (400) is 0.3-0.6 mm; the distance L between two adjacent spoiler structures (400) along the span of the blade is greater than D2+D3; 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 from the inlet end to the outlet end to adapt to the height change in the second chamber (200); the height of the spoiler structure (400) from the inlet end to the outlet end of the second chamber (200) is H1-Hn, H1 / F1 is in the range of 0.3-0.7, and Hn / F2 is in the range of 0.3-0.

7.

2. The turbine blade trailing edge cooling structure according to claim 1, characterized in that: The impingement air-entraining structure (300) comprises a plurality of impingement holes uniformly distributed along the radial direction of the blade.

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

0.

4. The turbine blade trailing edge cooling structure according to claim 1, characterized in that: The boss structure is cylindrical or hemispherical.

5. An aircraft engine, characterized in that: The turbine blade trailing edge cooling structure according to any one of claims 1 to 4 is applied.

Images