A turbine blade tip configuration and its design method

Abstract

This application discloses a turbine blade tip configuration and its design method, belonging to the technical field of aero-engine components. It includes film cooling holes (FSH), which allow cool air to be ejected from the FSH to form a film coating on the blade tip surface, thus cooling and protecting the heated surfaces of the turbine blade. A tip groove is formed at the tip of the turbine blade, and a shoulder wall is formed around the tip groove. Multiple shoulder wall grooves are spaced apart on the shoulder wall, dividing it into multiple shoulder wall protrusions. Multiple FSH holes are provided, each corresponding to a specific bottom of one of the shoulder wall grooves. These FSH holes allow cool air to be ejected along the tip grooves to cool the turbine blade tip. This application reduces the mass of the turbine blade through a lightweight turbine tip design, using a serrated groove on the shoulder wall of the grooved tip and arranging the FSH holes at the bottom of the serrated groove.

Description

Technical Field

[0001] This application relates to the field of aero-engine component technology, and in particular, to a turbine blade tip configuration. Furthermore, this application also relates to a design method applied to the aforementioned turbine blade tip configuration. Background Technology

[0002] The information provided in this section is for the purpose of generally presenting the background of this application. To the extent described in this section, the work of the currently named inventors and aspects of the description that may not constitute prior art at the time of filing are neither explicitly nor implicitly considered to be prior art of this application.

[0003] With the rapid development of aero engines, the requirements for their performance and economy are constantly increasing. Thrust-to-weight ratio and fuel efficiency are crucial indicators for evaluating aero engine performance and economy, and aero engines have extremely stringent requirements for these two metrics. To meet the needs of practical engineering, aero engines must achieve very high standards in terms of thrust-to-weight ratio and fuel efficiency, which often necessitates reducing overall mass.

[0004] As a crucial component of aero-engines, turbines require weight reduction designs to achieve overall engine weight reduction. On the other hand, with the continuous improvement of aero-engine performance parameters, combustion chamber exit temperatures are constantly rising, making the turbine's operating environment increasingly harsh. Currently, the turbine inlet temperature of advanced aero-engines far exceeds the limits that its blades can withstand. Therefore, while developing high-temperature resistant materials, designing efficient turbine cooling measures is particularly important. Ensuring the cooling performance of turbine blades while simultaneously reducing turbine weight under such harsh conditions is undoubtedly a challenging task.

[0005] To withstand the scouring of high-temperature combustion gases, turbine blades often employ a hollow internal structure. Furthermore, the turbine blade surface features numerous film cooling holes. During turbine operation, these holes eject cooling air to lower the surface temperature of the turbine blades, extending the turbine's lifespan. On the other hand, the hollow structure reduces turbine weight, offering a new approach to overall weight reduction. However, due to the unique operating conditions of turbines, excessive reduction in the internal structure is not advisable. This would cause the turbine blades, under high temperature and high speed conditions, to crack or even break due to the combined effects of centrifugal force and thermal stress, leading to serious accidents. For high-pressure turbines directly exposed to the high-temperature combustion gases at the combustion chamber outlet, the operating environment is even harsher. The minute leakage flow between the high-pressure turbine blades and the casing not only results in significant leakage flow losses but also generates a high heat transfer coefficient in the blade tip region, increasing the risk of turbine tip ablation. Therefore, reducing the mass of turbine blades while ensuring high tip cooling efficiency and aerodynamic performance becomes an extremely challenging problem.

[0006] In the prior art, such as the Chinese patent CN111379594A, a turbine rotor blade grooved tip structure and its design method are disclosed. This includes a top cover and a tip enclosure wall. The top cover is disposed on the top of the turbine rotor blade, and the tip enclosure wall is disposed around the top cover. The tip enclosure wall includes a pressure surface tip enclosure wall and a suction surface tip enclosure wall. The wall thickness L1 of the pressure surface tip enclosure wall is greater than the wall thickness L2 of the suction surface tip enclosure wall. The top cover and the tip enclosure wall form a tip groove. A variable radius rounded section is provided between the tip enclosure wall and the top cover. This application strengthens the tip enclosure wall itself, avoids excessively large tip cantilever dimensions, thereby increasing the strength margin at the corresponding location, alleviating stress concentration effects, thus improving blade lifespan, reducing replacement costs, and improving design reliability. This solution avoids stress concentration and improves the strength of the entire tip region by using a variable wall thickness design and a rounded section design for the tip enclosure wall, thereby enhancing reliability.

[0007] For example, Chinese Patent Publication No. CN112240228A discloses an intermittent groove for turbine blades with transverse slots, which discloses a blade tip structure for an intermittent groove with transverse slots on turbine blades. This structure includes a turbine blade, a blade tip, transverse slots, a suction-side groove wall, a pressure-side groove wall, and a cooling air passage. The blade tip is characterized by having a suction-side groove wall and a pressure-side groove wall, with transverse slots arranged at the fracture groove wall. The discontinuity of the intermittent groove is located on the pressure side, and the start of the discontinuity is located at the leading edge stagnation point. The suction-side groove wall and the pressure-side groove wall have the same height, the centerline of the transverse slot is parallel to the pressure surface arc, and the transverse slot communicates with the cooling air passage. This is achieved by setting a suction-side groove wall and a pressure-side groove wall on the blade tip, arranging transverse slots at the fracture groove wall, and ensuring the discontinuity of the intermittent groove is located on the pressure side and the start of the discontinuity is located at the leading edge stagnation point. By eliminating the gas re-attachment region formed within the intermittent groove at the blade tip, the heat transfer coefficient of the leading edge region is effectively reduced, the local heat transfer intensity is weakened, and the cooling characteristics of the blade tip are improved. The optimized length of the groove wall on the pressure side controls the lateral width of the groove, ensuring that the gas entering the groove forms only a single backflow vortex within the groove, eliminating the gas re-attachment region. The lateral slots effectively increase the lateral coverage of the jet after exiting the flow, improving the local film cooling efficiency at the blade tip, effectively increasing the working efficiency of the turbine blades, and enhancing the overall performance of the aero-engine.

[0008] The existing grooved blade tip design has been applied in the design of high-pressure turbine blades, as shown in the attached manual. Figure 5 As shown, due to the presence of the blade tip groove structure, the backflow vortex generated near the casing suppresses the gas leaking from the blade tip clearance, causing the leakage flow within the groove to repeatedly scour the bottom surface of the groove, thus increasing the heat transfer coefficient in these areas. However, as the groove depth increases, the leakage flow structure within the groove finds it difficult to directly scour the bottom of the groove, resulting in a decrease in the average heat transfer coefficient of the blade tip surface and poorer cooling at the blade tip. Reducing the shoulder wall thickness enhances the strength of the separation vortex within the groove, gradually increasing the heat transfer coefficient and thus benefiting blade tip cooling. However, due to strength limitations, excessively reducing the shoulder wall thickness is not feasible. In summary, arranging a groove structure at the blade tip can reduce the overall mass, but due to the influence of blade tip aerodynamic performance, heat transfer performance, and blade tip strength, the depth and area of ​​the groove are limited, and the mass reduction that can be achieved is also limited.

[0009] It should be noted that the information disclosed in the background section above is only used to enhance the understanding of the background of this application, and therefore may include information that does not constitute prior art known to those skilled in the art. Summary of the Invention

[0010] In view of at least one of the above technical problems, this application provides a turbine blade tip configuration that can reduce the mass of the turbine blade by a lightweight turbine blade tip design, which involves opening a serrated groove on the shoulder wall of the grooved blade tip and arranging the film cooling holes at the bottom of the serrated groove.

[0011] This application also provides a design method for the aforementioned turbine blade tip configuration.

[0012] According to one aspect of this application, a turbine blade tip configuration is provided, including a film cooling assembly disposed at the tip of the turbine blade. The film cooling assembly includes film cooling holes for spraying cool air from the film cooling holes to form a film covering layer on the tip surface to cool and protect the heated wall surface of the turbine blade. The turbine blade tip configuration also includes a tip groove, a shoulder wall, and a shoulder wall groove.

[0013] The blade tip groove is formed at the tip of the turbine blade. A shoulder wall is formed around the blade tip groove at the tip of the turbine blade. Multiple shoulder wall grooves are spaced apart on the shoulder wall, and the multiple shoulder wall grooves divide the shoulder wall to form multiple shoulder wall protrusions. Multiple film cooling holes are formed, and the multiple film cooling holes are correspondingly formed at the bottom of the multiple shoulder wall grooves. The film cooling holes are used to spray cold air into the blade tip groove to cool the tip of the turbine blade.

[0014] In some embodiments of this application, the outer contour curve of the blade tip groove is adapted to the airfoil curve of the turbine blade, and the depth of the blade tip groove is consistent with the depth of the shoulder wall groove.

[0015] In some embodiments of this application, the depth of both the blade tip groove and the shoulder wall groove is no greater than 2 mm.

[0016] In some embodiments of this application, the side of the shoulder wall is perpendicular to the tip surface of the turbine blade, and the thickness of the shoulder wall is not less than 0.5 mm.

[0017] In some embodiments of this application, the film air holes are oblique holes, and each film air hole is inclined toward the blade tip groove.

[0018] In some embodiments of this application, the width of the shoulder wall protrusion at the trailing edge of the turbine blade is 1.1-1.2 times the width of the shoulder wall groove.

[0019] According to another aspect of this application, a turbine blade tip configuration design method is also provided, which is applied to the aforementioned turbine blade tip configuration. The turbine blade tip configuration design method includes the following steps:

[0020] S100, The outer contour curve of the blade tip groove is formed by offsetting and reducing the blade profile curve;

[0021] S200. Grooves are made on the tip surface of the turbine blade according to the outer contour curve of the tip groove to form the tip groove, and a shoulder wall is formed on the outer ring of the tip groove.

[0022] S300: Multiple shoulder wall grooves are opened at intervals on the shoulder wall to divide the shoulder wall into multiple shoulder wall protrusions;

[0023] S400, air film holes are opened at the bottom of each shoulder wall groove.

[0024] In some embodiments of this application, the offset of the blade profile curve in step S100 is the thickness of the shoulder wall, and the offset of the blade profile curve is not less than 0.5 mm.

[0025] In some embodiments of this application, in step S300, except for the shoulder wall groove at the trailing edge of the turbine blade, the width of each shoulder wall groove is the same.

[0026] In some embodiments of this application, the drilling angle of the air film hole in step S400 is inclined toward the blade tip groove, and the inclination angle is 18°-22°.

[0027] This application has the following beneficial effects:

[0028] This application discloses a turbine blade tip configuration that reduces weight by creating a tip groove at the blade tip. Furthermore, multiple intermittent shoulder wall grooves on the outer ring of the tip groove further reduce the overall mass of the turbine blade. Film cooling holes are created at the bottom of the shoulder wall grooves, allowing the cooling gas ejected from these holes to adhere more closely to the surface of the tip groove, preventing ablation at the turbine blade tip. Moreover, arranging the film cooling holes at the bottom of the shoulder wall grooves, compared to creating holes at the top of the shoulder wall or other locations, allows for easier positioning based on the individual shoulder wall grooves, reducing manufacturing complexity.

[0029] The turbine blade tip configuration design method of this application also has the aforementioned beneficial effects. It also includes offsetting and reducing the outer contour of the tip groove based on the blade profile curve of the turbine blade, eliminating the need for additional design of the outer contour curve of the tip groove, which helps to reduce design costs and significantly reduces the difficulty of processing and inspection. In addition, it can also achieve consistency in the thickness of the shoulder wall protrusions at various locations, ensuring overall strength and effectively reducing the overall mass of the turbine blade.

[0030] Of course, any product implementing this application does not necessarily need to achieve all the advantages described above simultaneously. In addition to the purposes, features, and advantages described above, this application also has other purposes, features, and advantages. The following will provide a more detailed description of this application with reference to figures. Attached Figure Description

[0031] The accompanying drawings, which form part of this application, are used to provide a further understanding of this application. The illustrative embodiments and descriptions of this application are used to explain this application and do not constitute an undue limitation of this application. In the drawings:

[0032] Figure 1 This is a schematic diagram of the overall structure of a preferred embodiment of this application;

[0033] Figure 2 This is a schematic diagram of the shoulder wall protrusion and shoulder wall groove of a preferred embodiment of this application;

[0034] Figure 3 This is a schematic diagram of the blade tip groove according to a preferred embodiment of this application;

[0035] Figure 4 This is a schematic diagram showing the location of the air film pores in a preferred embodiment of this application;

[0036] Figure 5 This is a schematic diagram of a turbine blade with a groove at the tip, based on existing technology.

[0037] Legend: 100, turbine blade; 200, casing; 1, blade tip groove; 2, shoulder wall; 21, shoulder wall protrusion; 3, shoulder wall groove; 4, film gas hole. Detailed Implementation

[0038] The embodiments of this application are described in detail below with reference to the accompanying drawings; however, this application may be implemented in a variety of different ways as defined and covered below.

[0039] A turbine blade tip configuration includes a film cooling assembly disposed at the tip of a turbine blade 100. The film cooling assembly includes film cooling holes 4, which are used to spray cool air from the film cooling holes 4 to form a film covering layer on the tip surface of the blade, thereby cooling and protecting the heated wall surface of the turbine blade 100. The turbine blade tip configuration also includes a tip groove 1, a shoulder wall 2, and a shoulder wall groove 3.

[0040] A blade tip groove 1 is formed at the blade tip of the turbine blade 100. A shoulder wall 2 is formed around the blade tip groove 1 at the blade tip of the turbine blade 100. Multiple shoulder wall grooves 3 are spaced apart on the shoulder wall 2. The multiple shoulder wall grooves 3 divide the shoulder wall 2 to form multiple shoulder wall protrusions 21. Multiple film cooling holes 4 are formed. The multiple film cooling holes 4 are formed one-to-one at the bottom of the multiple shoulder wall grooves 3. The film cooling holes 4 are used to spray cold air into the blade tip groove 1 to cool the blade tip of the turbine blade 100.

[0041] Here, "shoulder wall 2" refers to a ring of sidewall-like structures around the outer edge of the blade tip groove 1. In some embodiments, shoulder wall 2 is a ring of plate-like structures with equal thickness and height. The inner contour of shoulder wall 2 is the outer contour of blade tip groove 1. Both the inner and outer contours of shoulder wall 2 are consistent with the blade profile curve of turbine blade 100.

[0042] The cold air in the air mold hole 4 can be cooling gas drawn from the compressor or other external air sources, depending on the specific situation.

[0043] The turbine blade tip configuration of this application reduces weight by creating a tip groove 1 at the tip of the turbine blade 100. Simultaneously, multiple shoulder wall grooves 3 are intermittently created on the shoulder wall 2 surrounding the tip groove 1, further reducing the overall mass of the turbine blade 100. Film cooling holes 4 are created at the bottom of the shoulder wall grooves 3, allowing the cooling gas ejected from the holes 4 to adhere more closely to the surface of the tip groove 1, preventing easy ablation of the turbine blade tip. Furthermore, arranging the film cooling holes 4 at the bottom of the shoulder wall grooves 3, compared to creating holes at the top of the shoulder wall 2 or other locations, allows for easier positioning based on the individual shoulder wall grooves 3 and reduces manufacturing difficulty.

[0044] Preferably, please refer to Figure 1 and 3 As shown, the outer contour curve of the blade tip groove 1 is adapted to the airfoil curve of the turbine blade 100, and the depth of the blade tip groove 1 is consistent with the depth of the shoulder wall groove 3.

[0045] It is understandable that the outer contour of the blade tip groove 1 is obtained by offsetting and reducing the airfoil curve of the turbine blade 100, which can reduce the design cost of the blade tip groove 1 and facilitate processing. This can also achieve the consistency of the thickness of the shoulder wall 2 in all places and ensure its overall strength.

[0046] Optionally, the outer contour curve of the tip groove 1 is formed by an approximate curve of the turbine blade tip shoulder wall thickness required by the turbine blade profile offset. In some embodiments, the outer contour curve of the tip groove 1 adopts a cubic Bezier curve controlled by four points, generating a smooth curve based on the start point, end point, and two intermediate control points. The outer contour curve of the tip groove 1, i.e., the thickness distribution of the tip groove 1, can also be determined by the thickness distribution of the turbine blade profile 100 and the predetermined thickness of the shoulder wall 2.

[0047] In this preferred embodiment, the depth of both the blade tip groove 1 and the shoulder wall groove 3 is no greater than 2 mm.

[0048] It should be noted that the depth of the blade tip groove 1 is consistent with the depth of the shoulder wall groove 3. This reduces the machining difficulty of the shoulder wall groove 3 and also effectively reduces the positioning and machining difficulty of the film cooling hole 4 at the bottom of the shoulder wall groove 3. However, the depth of the blade tip groove 1 should not be too deep. An excessively deep groove 1 will increase the fluid flow within it, reducing the chance of direct contact between the leakage flow and the bottom of the groove 1, thus lowering the heat transfer coefficient. The depth of the blade tip groove 1 should be controlled to be no greater than 2mm, because when the groove depth exceeds 2mm, the heat transfer coefficient decreases significantly. Therefore, controlling the depth of both the blade tip groove 1 and the shoulder wall groove 3 to be no greater than 2mm effectively ensures the heat transfer coefficient of the blade tip groove 1, ensuring the heat transfer efficiency of the turbine blade 100 tip. Combined with the film cooling hole 4 spraying cooling gas towards the blade tip groove 1, this achieves a good blade tip cooling and protection effect.

[0049] Preferably, please refer to Figure 1 and 4 As shown, the side of the shoulder wall 2 is perpendicular to the tip surface of the turbine blade 100, and the thickness of the shoulder wall 2 is not less than 0.5 mm.

[0050] The thickness of the shoulder wall 2 here refers to the distance from the inner side of the shoulder wall 2 near the tip groove 1 to the outer side of the shoulder wall 2 flush with the side wall of the turbine blade 100 when the tip surface of the turbine blade 100 is horizontal. The side of the shoulder wall 2 is perpendicular to the tip surface of the turbine blade 100, which means that the side of the tip groove 1 structure has no tilt angle with the tip surface of the turbine blade 100, and ensures that the side of the tip groove 1 is perpendicular to the bottom surface of the tip groove 1. This helps to reduce the difficulty of processing and ensures smooth heat exchange in the tip groove 1.

[0051] Understandably, given a fixed overall area at the turbine blade tip (100mm), the thickness of the shoulder wall 2 affects the area of ​​the tip groove 1. It should be noted that the area of ​​the tip groove 1 should be controlled. If the area of ​​the tip groove 1 is too large, the separation vortex will fully develop within it, increasing its intensity and thus the heat transfer coefficient on the surface of the tip groove 1. However, due to the limitations of the turbine's special operating environment, an excessively large tip groove 1 area can easily lead to a thinner shoulder wall 2, resulting in reduced strength. Therefore, the thickness of the shoulder wall 2 should be controlled to be no less than 0.5mm to ensure sufficient strength and prevent impact damage.

[0052] Optionally, the thickness of the shoulder wall 2 is up to 0.9 mm, which can avoid encroaching too much on the area of ​​the blade tip groove 1 and affecting the heat transfer coefficient of the blade tip groove 1, while also not increasing the overall mass of the turbine blade 100 too much, thus ensuring the achievement of the weight reduction goal.

[0053] Preferably, please refer to Figure 1 As shown, the air film pore 4 is an oblique hole structure, and each air film pore 4 is inclined toward the blade tip groove 1.

[0054] It is understandable that the air film holes 4 are all inclined toward the blade tip groove 1, which can quickly spray the cooling gas into the blade tip groove 1, so as to achieve the coverage of the cooling gas 1 on the blade tip of the turbine blade 100, and achieve a stable and effective cooling effect.

[0055] Preferably, please refer to Figure 3 and 4 As shown, the width of the shoulder wall protrusion 21 at the trailing edge of the turbine blade 100 is 1.1-1.2 times the width of the shoulder wall groove 3.

[0056] The width of the shoulder wall protrusion 21 here refers to the distance between adjacent shoulder wall grooves 3 below the horizontal plane of the turbine blade tip 100. Similarly, the width of the shoulder wall groove 3 refers to the horizontal distance between adjacent shoulder wall protrusions 21.

[0057] It should be noted that the overall thickness at the trailing edge is relatively small. Considering the strength of the trailing edge, it is not advisable to open too many or too wide shoulder wall grooves 3 near the trailing edge. By adjusting the width of the shoulder wall grooves 3, the shoulder wall protrusion 21 at the trailing edge is 1.1-1.2 times the width of the shoulder wall grooves 3. That is, reducing the width of the shoulder wall grooves 3 at the trailing edge can effectively optimize the blade tip structure at the trailing edge and ensure the strength of the trailing edge.

[0058] According to another aspect of this application, a turbine blade tip configuration design method is also provided, which is applied to the aforementioned turbine blade tip configuration. The turbine blade tip configuration design method includes the following steps:

[0059] S100, The outer contour curve of the tip groove 1 is formed by offsetting and reducing the airfoil curve of the turbine blade 100;

[0060] S200. Grooves are made on the tip surface of the turbine blade 100 according to the outer contour curve of the tip groove 1 to form the tip groove 1, and a shoulder wall 2 is formed on the outer ring of the tip groove 1.

[0061] S300. Multiple shoulder wall grooves 3 are opened at intervals on the shoulder wall 2 to divide the shoulder wall 2 into multiple shoulder wall protrusions 21.

[0062] S400, air film holes 4 are respectively opened at the bottom of each shoulder wall groove 3.

[0063] In some embodiments of this application, the width of the shoulder wall groove 3 is d1, the width of the shoulder wall protrusion 21 is d2, the diameter of the film cooling hole 4 is d3, and the distance between the top of the shoulder wall protrusion 21 and the casing 200 is the tip clearance d4. d1 is between 1 and 1.5 mm, which effectively ensures the heat transfer coefficient at the tip groove 1 and the overall strength of the shoulder wall 2. However, d1 should not be too small, so as not to affect the cooling outflow distance and cooling effect of the film cooling hole 4 at the bottom of the shoulder wall groove 3. After determining the width of the shoulder wall groove 3, grooves can be made on the shoulder wall 2 to form various shoulder wall protrusions 21. The diameter d3 of the film cooling hole 4 is 0.2-0.4 mm, which can achieve efficient and stable injection of cooling gas. d4 is a predetermined design value, which is 1.0%-1.5% of the turbine blade height.

[0064] The turbine blade tip configuration design method of this application also has the aforementioned beneficial effects. It also includes offsetting and reducing the outer contour of the tip groove 1 based on the airfoil curve of the turbine blade 100, eliminating the need for additional design of the outer contour curve of the tip groove, which helps to reduce design costs and significantly reduces the difficulty of processing and inspection. In addition, it can also achieve the consistency of the thickness of the shoulder wall protrusions 21 in various places, ensuring the overall strength, while effectively reducing the overall mass of the turbine blade.

[0065] Preferably, in step S100, the offset of the blade profile curve of the turbine blade 100 is the thickness of the shoulder wall 2, and the offset of the blade profile curve is not less than 0.5 mm.

[0066] Understandably, the outer contour curve of the blade tip groove 1 is consistent with the inner and outer curves of the shoulder wall 2, and both are consistent with the airfoil curve of the turbine blade 100. This reduces the design cost of the contour curves and lowers the difficulty of processing and inspection. However, if the area of ​​the blade tip groove 1 is too large, the shoulder wall 2 will be too thin, resulting in a decrease in the strength of the shoulder wall 2. Therefore, the thickness of the shoulder wall 2 is controlled to be no less than 0.5 mm to ensure that the shoulder wall 2 has sufficient strength and is not prone to impact damage. At the same time, the offset of the airfoil curve, i.e., the maximum thickness of the shoulder wall 2, is 0.9 mm. This avoids encroaching too much on the area of ​​the blade tip groove 1, which would affect the heat transfer coefficient of the blade tip groove 1, and also does not excessively increase the overall mass of the turbine blade 100, thus ensuring that the weight reduction goal is achieved.

[0067] Preferably, in step S300, except for the shoulder wall groove at the trailing edge of the turbine blade 100, the width of each shoulder wall groove 3 is the same.

[0068] It is understandable that the thickness at the trailing edge of the turbine blade 100 is relatively small. In order to ensure the strength of the trailing edge, too many or too wide shoulder wall grooves 3 should be avoided near the trailing edge. By making the shoulder wall protrusion 21 at the trailing edge 1.1-1.2 times the width of the shoulder wall groove 3, that is, by reducing the width of the shoulder wall groove 3 at the trailing edge, the blade tip structure at the trailing edge can be effectively optimized and the strength of the trailing edge can be guaranteed.

[0069] Preferably, in step S400, the drilling angle of the air film hole 4 is inclined toward the blade tip groove 1, and the inclination angle is 18°-22°.

[0070] Understandably, by tilting the film cooling orifice 4 toward the blade tip groove 1, cooling gas can be smoothly sprayed toward the blade tip groove 1, effectively covering the tip of the turbine blade 100 and ensuring the cooling protection effect. However, the tilt angle of the film cooling orifice 4 should not be too large, as this will easily cause the cooling airflow to dissipate, which is not conducive to forming film coverage. At the same time, the tilt angle should not be too small, as this will easily lead to problems such as incomplete cooling gas coverage. Therefore, the tilt angle of the film cooling orifice 4 is controlled at 18°-22° to achieve the optimal range of cooling gas injection and coverage.

[0071] In summary, the main objective of this application is to address the weight reduction problem of high-pressure turbine blades by designing the blade tip structure. Conventional weight reduction methods cannot meet the needs of existing practical engineering projects, therefore, a lightweight turbine blade tip design method is proposed. A serrated shoulder wall groove 3 is formed on the shoulder wall 2 of the turbine blade 100 with a blade tip groove 1, further reducing the overall mass of the turbine blade 100. Film cooling holes 4 are arranged at the bottom of the serrated shoulder wall groove 3 to enhance heat transfer at the bottom of the blade tip groove 1, allowing the cooling gas ejected from the film cooling holes 4 to adhere more closely to the surface of the blade tip groove 1, preventing blade tip ablation. The film cooling holes 4 are easy to position and their machining difficulty is effectively reduced when arranged at the bottom of the shoulder wall groove 3. High-pressure turbine blades manufactured using the design method of this application are characterized by light overall weight and strong blade tip cooling gas coverage. This application largely meets the needs of existing engineering projects for weight reduction of high-pressure turbine blade tips.

[0072] It should be noted that, in this document, the terms “comprising,” “including,” or any other variations thereof are intended to cover non-exclusive inclusion, such that a process, method, article, or apparatus that comprises a list of elements includes not only those elements but also other elements not expressly listed, or elements inherent to such process, method, article, or apparatus.

[0073] This document uses specific examples to illustrate the principles and implementation methods of this application. The examples are merely for the purpose of helping to understand the methods and core ideas of this application. The above descriptions are only preferred embodiments of this application. It should be noted that due to the limitations of written expression, and the existence of an infinite number of specific structures, those skilled in the art can make various improvements, modifications, or variations without departing from the principles of this application, and can also combine the above technical features in an appropriate manner. These improvements, modifications, variations, or combinations, or the direct application of the inventive concept and technical solution to other situations without modification, should all be considered as protected by this application.

Claims

1. A turbine blade tip configuration, comprising a film cooling assembly disposed at the tip of a turbine blade (100), the film cooling assembly comprising film cooling holes (4), the film cooling holes (4) being used to eject cool air from the film cooling holes (4) to form a film covering layer on the tip surface of the blade, thereby cooling and protecting the heated wall surface of the turbine blade (100), characterized in that, The turbine blade tip configuration also includes a tip groove (1), a shoulder wall (2), and a shoulder wall groove (3): A blade tip groove (1) is formed on the blade tip of the turbine blade (100). A shoulder wall (2) is formed around the blade tip groove (1) on the blade tip of the turbine blade (100). Multiple shoulder wall grooves (3) are spaced apart on the shoulder wall (2). The multiple shoulder wall grooves (3) divide the shoulder wall (2) to form multiple shoulder wall protrusions (21). Multiple air film holes (4) are formed. The multiple air film holes (4) are formed one-to-one at the bottom of the multiple shoulder wall grooves (3). The air film holes (4) are used to spray cold air into the blade tip groove (1) to cool the blade tip of the turbine blade (100).

2. The turbine blade tip configuration according to claim 1, characterized in that, The outer contour curve of the blade tip groove (1) is adapted to the blade profile curve of the turbine blade (100), and the depth of the blade tip groove (1) is consistent with the depth of the shoulder wall groove (3).

3. A turbine blade tip configuration according to claim 1 or 2, characterized in that, The depth of the blade tip groove (1) and the shoulder wall groove (3) is no greater than 2 mm.

4. The turbine blade tip configuration according to claim 1, characterized in that, The side of the shoulder wall (2) is perpendicular to the tip surface of the turbine blade (100), and the thickness of the shoulder wall (2) is not less than 0.5 mm.

5. The turbine blade tip configuration according to claim 1, characterized in that, The air film pores (4) are oblique holes, and each air film pore (4) is inclined toward the blade tip groove (1).

6. The turbine blade tip configuration according to claim 1, characterized in that, The width of the shoulder wall protrusion (21) at the tip trailing edge of the turbine blade (100) is 1.1-1.2 times the width of the shoulder wall groove (3).

7. A method for designing the tip configuration of a turbine blade, characterized in that, Applied to the turbine blade tip configuration as described in any one of claims 1-6, the turbine blade tip configuration design method includes the following steps: S100, The outer contour curve of the tip groove (1) is formed by offsetting and reducing the airfoil curve of the turbine blade (100); S200. Grooves are made on the tip surface of the turbine blade (100) according to the outer contour curve of the tip groove (1) to form the tip groove (1), and a shoulder wall (2) is formed on the outer ring of the tip groove (1); S300. Multiple shoulder wall grooves (3) are opened at intervals on the shoulder wall (2) to separate the shoulder wall (2) and form multiple shoulder wall protrusions (21). S400, air film holes (4) are opened at the bottom of each shoulder wall groove (3).

8. The turbine blade tip configuration design method according to claim 7, characterized in that, In step S100, the offset of the blade profile curve of the turbine blade (100) is the thickness of the shoulder wall (2), and the offset of the blade profile curve is not less than 0.5 mm.

9. The turbine blade tip configuration design method according to claim 7, characterized in that, In step S300, except for the shoulder wall groove at the trailing edge of the turbine blade (100), the width of each shoulder wall groove (3) is the same.

10. The turbine blade tip configuration design method according to claim 7, characterized in that, In step S400, the drilling angle of the air film hole (4) is inclined toward the blade tip groove (1), and the inclination angle is 18°-22°.

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