Compound tip structure and blade for improving gas thermal performance of turbine rotor tip

Abstract

The application discloses a compound blade tip structure and a blade for improving the gas-heat performance of a turbine rotor tip, wherein a groove for restraining tip leakage flow is arranged on the end face of the tip, and a certain number of staggered semicircular serrations are arranged along the flow direction from the leading edge to the trailing edge in the groove; the staggered semicircular serration structure changes the flow state of the leakage flow after entering the groove, blocks the development of the cavity vortex and other groove vortex systems, effectively restrains the gap leakage flow, and achieves the purpose of improving the turbine aerodynamic efficiency; on the other hand, the arrangement position of the semicircular serration structure changes the high heat exchange degree area formed by the high-temperature gas impacting the tip surface, reduces the heat exchange coefficient of the tip surface, improves the heat exchange characteristics of the tip area, and improves the overall performance of the aero-engine.

Description

Technical Field

[0001] This invention relates to the field of gas turbine blade cooling technology, specifically to a composite blade tip structure and blade for improving the aerothermal performance of turbine rotor blade tips. Background Technology

[0002] Turbines are key components of power plants such as aero engines and ground-based gas turbines, widely used in energy, defense, and other fields, and are important strategic technological equipment. For a long time, turbine rotor tip leakage and the resulting losses have been a significant aerodynamic problem attracting widespread attention from researchers. Research data shows that tip leakage losses in high-pressure turbines account for approximately one-third of the total losses within the turbine's flow path. Furthermore, because the airflow does no work on the turbine, the high-temperature leakage flow leads to high temperatures and thermal loads on the blade tip walls, thus threatening the reliability and lifespan of the turbine tip region. How to effectively reduce tip leakage losses and utilize flow control to reduce the thermal load on the blade tip is one of the crucial issues that must be addressed in the design of high-efficiency, high-reliability high-pressure turbines.

[0003] Tip leakage flow control methods can be divided into active and passive control methods. Active control methods include plasma excitation technology and active tip clearance control systems. Although active control methods can effectively block tip leakage flow, the required control systems may introduce additional problems, thus limiting their practical engineering applications. Currently, the most commonly used tip leakage flow control method in aero-engines is passive control, with tip modification as a typical measure. Tip structural designs include ribbed tips, winglet tips, and hybrid designs. Researchers have long conducted extensive studies on the aerothermal performance of different tip structures, finding that the tip geometry significantly influences its aerothermal performance. Summary of the Invention

[0004] To address the problems existing in the prior art, the present invention provides a composite blade tip structure and blade that improves the aero-thermal performance of turbine rotor blade tips. This structure alters the flow structure inside the blade tip groove, affecting the scouring and reattachment positions of the blade tip leakage flow on the groove surface, reducing the blade tip leakage flow, lowering the heat transfer coefficient in the local area of ​​the blade tip, and improving the overall aero-thermal characteristics of the grooved blade tip.

[0005] This invention is achieved through the following technical solution:

[0006] A composite blade tip structure for improving the aero-thermal performance of a turbine rotor blade tip includes grates and grooves for suppressing tip leakage flow. The grooves are disposed on the tip face of the blade, and multiple grates are disposed on the tip face and located in the grooves. The multiple grates are arranged at intervals along the leading edge to the trailing edge of the blade. The transverse direction of the grates is perpendicular to the mid-arc line of the blade, and the top surface of the grates is arc-shaped and arranged transversely.

[0007] Preferably, the plurality of grating teeth are arranged alternately along the central arc direction.

[0008] Preferably, the teeth are semi-circular.

[0009] Preferably, the height of the comb teeth is 0.5 to 1.5 times the height of the blade tip gap.

[0010] Preferably, the connection between the top surface of the comb tooth and the side wall is an outer arc surface, and the connection between the side wall of the comb tooth and the tip surface of the blade is an inner arc surface.

[0011] Preferably, the radius of the outer arc surface is 0.1 to 0.3 mm, and the radius of the inner arc surface is 0.2 to 0.5 mm.

[0012] Preferably, the sidewalls of the comb teeth are inclined surfaces with an inclination angle α of 10° to 30°.

[0013] Preferably, the distance D between the starting tooth and the leading edge of the blade is 0.1 to 0.2 times the chord length C of the blade; the spacing E between two adjacent teeth is 0.1 to 0.2 times the chord length C of the blade.

[0014] Preferably, the depth of the groove is 0.5 to 2.0 times the tip clearance height G.

[0015] A turbine blade having the aforementioned composite tip structure at its tip.

[0016] Compared with the prior art, the present invention has the following beneficial technical effects:

[0017] This invention proposes a composite blade tip structure to improve the aero-thermal performance of turbine rotor blade tips. By incorporating a groove structure on the end face of the blade tip and setting multiple semi-circular grates within the groove, these grates are spaced apart along the mid-arc line, with their transverse direction perpendicular to the mid-arc line. This staggered arrangement of semi-circular grates alters the area of ​​leakage gas impact and re-adhesion within the grooved blade tip, effectively reducing the heat transfer intensity on the near-suction side of the blade tip leading edge. Furthermore, the leakage flow is impeded by the rounded-corner near-pressure side semi-circular grates as it passes through the blade tip gap, thereby reducing blade tip leakage. The rational staggered layout, placement, and geometry of the semi-circular grates alter the aero-thermal characteristics of the grooved blade tip, effectively improving the turbine rotor's operating efficiency and the overall performance of the aero-engine. This invention optimizes the blade tip structure to reduce leakage losses, thereby improving aerodynamic performance and ultimately enhancing the heat transfer characteristics of the blade tip region. Attached Figure Description

[0018] Figure 1 This is a schematic diagram of the composite grooved blade tip structure of the semi-circular comb teeth of the present invention;

[0019] Figure 2 This is a top view of the composite grooved blade tip structure of the semi-circular comb teeth of the present invention;

[0020] Figure 3 This is a cross-sectional view of the composite grooved blade tip structure of the semi-circular comb teeth of the present invention;

[0021] Figure 4 for Figure 3 Enlarged view of the structure of the central comb teeth;

[0022] Figure 5 This is a cloud diagram of the tip heat transfer coefficient of the composite grooved blade tip structure with semi-circular grating teeth of the present invention.

[0023] Figure 6 This is a schematic diagram of the partial circumferential cross-sectional streamline distribution of the composite groove tip structure of the semi-circular comb teeth of the present invention.

[0024] In the diagram: 1-blade, 2-blade tip, 3-suction surface shoulder wall, 4-semi-circular serrations, 5-pressure surface shoulder wall. Detailed Implementation

[0025] The present invention will now be described in further detail with reference to the accompanying drawings. These descriptions are intended to explain the invention and not to limit it.

[0026] See Figure 1-6 A composite blade tip structure for improving the aero-thermal performance of turbine rotor blade tips includes grates and grooves for changing leakage flow. The grooves are disposed on the blade tip face, and multiple grates are disposed on the blade tip face and located in the grooves. The multiple grates are arranged at intervals along the mid-arc line of the blade from the leading edge to the trailing edge. The transverse direction of the grates is perpendicular to the mid-arc line of the blade, and the top surface of the grates is arc-shaped and arranged in the transverse direction.

[0027] This composite grooved blade tip structure features grooves on the blade tip surface, altering the flow state of the leakage flow after it enters the grooves. This blocks the development of grooved vortex systems such as chamber vortices, effectively suppressing gap leakage flow and improving turbine aerodynamic efficiency. Secondly, the arrangement of grates in the grooves changes the high heat transfer area formed by the impact of high-temperature combustion gas on the blade tip surface, reducing the heat transfer coefficient of the blade tip surface, improving the heat transfer characteristics of the blade tip region, and enhancing the overall performance of the aero-engine.

[0028] The blade tip has an upwardly protruding, circumferentially closed shoulder wall at its edge. The area formed by the inner wall of the shoulder wall and the tip of the blade is a groove. The shoulder wall includes a suction surface shoulder wall 3 and a pressure surface shoulder wall 5, which are connected end to end to form a circumferentially closed structure. The suction surface shoulder wall 3 and the pressure surface shoulder wall 5 have the same width W and height H.

[0029] The distance from the top surface of the shoulder wall to the casing is the blade tip clearance G, the width W of the shoulder wall is 0.6 to 1.2 times the blade tip clearance G, and the groove depth H is 0.5 to 2.0 times the blade tip clearance height G.

[0030] The multiple grating teeth are arranged alternately along the mid-arc line. For example, starting from the blade edge, the first grating tooth is set on the pressure surface side of the mid-arc line, the second grating tooth is set on the suction surface side of the mid-arc line, and so on. The distance D between the first grating tooth and the leading edge of the blade is 0.1 to 0.2 times the blade chord length C. The spacing E between two adjacent grating teeth along the mid-arc line is 0.1 to 0.2 times the blade chord length C. The blade tip gap height G and the blade chord length C are 1 mm and 57 mm, respectively.

[0031] The grating teeth have a semi-circular structure, with a height of L, which is the radius of the grating teeth. The distance from the first grating tooth to the leading edge of the blade is D, and the distance between two adjacent grating teeth along the central arc is E.

[0032] The height L of the grating teeth is 0.5 to 1.5 times the height of the blade tip gap; the radius R1 of the rounded corner between the top surface and the side wall of the grating teeth is 0.1 to 0.3 mm; the radius R2 of the rounded corner between the tip surface of the grating teeth is 0.2 to 0.5 mm; and the slope α of the side wall of the grating teeth is 10° to 30°.

[0033] The number of semi-circular grates is 3 to 6, preferably 4. The four grates are arranged laterally perpendicular to the mid-arc line of the blade tip, and are staggered along the mid-arc line. This arrangement of multiple grates alters the fluid flow state after the leakage flow from the blade tip enters the groove under the action of the lateral pressure difference on both sides of the blade. The grates near the pressure side mainly hinder the development of the chamber vortex, causing the leakage flow to form a backflow vortex after entering the groove, reducing the amount of leakage at the blade tip and lowering the aerodynamic losses in the blade tip region. The grates near the suction side change the gas scouring and re-attachment area formed in the grooved blade tip, effectively reducing the heat transfer coefficient in the near-suction side region and weakening the local heat transfer intensity. The staggered arrangement and geometry of the grates effectively change the aero-thermal characteristics of the grooved blade tip, extending the working life of the blade.

[0034] Example 1

[0035] A composite blade tip structure for improving the aero-thermal performance of turbine rotor blade tips is provided with suction surface shoulder wall 3 and pressure surface shoulder wall 5 of equal height and width on the blade tip 2, forming a groove structure with the blade tip tip face. Semi-circular grating teeth 4 are arranged at a certain distance from the leading edge of the blade. The transverse direction of multiple grating teeth 4 is perpendicular to the middle arc line of the blade tip and is staggered along the middle arc line.

[0036] The groove shoulder wall width W is 1.0 mm; the groove depth H is 1.5 mm; the grate tooth height L is 1.5 mm; the upper fillet radius R1 of the grate tooth is 0.3 mm; the lower fillet radius R2 of the grate tooth is 0.2 mm; the grate tooth sidewall inclination angle α is 10°; the distance D from the starting position of the first grate tooth to the leading edge of the blade is 0.1 C; the distance E between two adjacent grate teeth is 0.15 C; the blade tip gap height G and the blade chord length C are 1 mm and 57 mm, respectively.

[0037] See Figure 5 and 6 In this embodiment, a groove is provided on the blade tip surface. Within the groove are three major vortices: a scraping vortex, a pressure-side vortex, and a suction-side rib corner vortex. These three interact, altering the aerodynamic and thermodynamic layout of the blade tip, weakening or even eliminating the blade tip gap vortex intensity, and consequently changing the vortex structure within the groove. The scraping vortex within the groove acts as an aerodynamic grate, effectively reducing leakage and the normal momentum of the leakage flow, thereby reducing external mixing losses. The height of the grate in the groove is the same as the depth of the groove shoulder wall. By reducing the leakage flow area within the groove and hindering the development of the chamber vortex, a more complex vortex structure is formed within the groove, effectively reducing blade tip leakage and aerodynamic losses caused by blade tip leakage. Simultaneously, the upper and lower rounded corners of the grate have a stabilizing effect on the leaking combustion gas, changing the impact area of ​​the leakage flow and thus improving the heat transfer characteristics of the blade tip surface.

[0038] In another embodiment, a blade is also provided, wherein the blade tip surface is provided with the above-mentioned composite grooved blade tip structure for improving the aero-thermal performance of turbine rotor blade tips.

[0039] The above content is only for illustrating the technical concept of the present invention and should not be construed as limiting the scope of protection of the present invention. Any modifications made to the technical solution based on the technical concept proposed in this invention shall fall within the scope of protection of the claims of this invention.

Claims

1. A composite blade tip structure for improving the aero-thermal performance of turbine rotor blade tips, characterized in that, It includes grating teeth and grooves for suppressing leakage flow at the blade tip. The grooves are set on the blade tip face, and multiple grating teeth are set on the blade tip face and located in the grooves. The multiple grating teeth are arranged at intervals along the leading edge to the trailing edge of the blade. The transverse direction of the grating teeth is perpendicular to the mid-arc line of the blade, and the top surface of the grating teeth is arc-shaped and set in the transverse direction. The height of the grating teeth is 0.5 to 1.5 times the height of the blade tip gap; The sidewalls of the comb teeth are inclined surfaces with an inclination angle α of 10° to 30°. The distance D between the starting teeth and the leading edge of the blade is 0.1 to 0.2 times the chord length C of the blade; the spacing E between two adjacent teeth is 0.1 to 0.2 times the chord length C of the blade. The depth of the groove is 0.5 to 2.0 times the tip clearance height G; The multiple sieve teeth are arranged in an alternating pattern along the central arc direction; The connection between the top surface of the comb tooth and the side wall is an outer arc surface, and the connection between the side wall of the comb tooth and the tip surface of the blade is an inner arc surface.

2. The composite blade tip structure for improving the aero-thermal performance of a turbine rotor blade tip according to claim 1, characterized in that, The comb teeth have a semi-circular structure.

3. The composite blade tip structure for improving the aero-thermal performance of a turbine rotor blade tip according to claim 1, characterized in that, The radius of the outer arc surface is 0.1~0.3mm, and the radius of the inner arc surface is 0.2~0.5mm.

4. A turbine blade, characterized in that, The blade tip is provided with the composite blade tip structure as described in any one of claims 1-3.

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