A method for inhibiting local failure of a turbine blade trailing edge film cooling

Abstract

A method for suppressing localized film gas failure at the trailing edge of a turbine blade includes: dividing the trailing half of the blade root into a film gas failure triangle zone, and adding a non-uniform array of film gas holes within the zone; the length of the film gas failure triangle zone is 80% of the axial chord length from the trailing edge of the blade, and the height is 35% of the blade height from the blade root; the lateral spacing of the film gas holes is set according to the position of the longitudinal air chamber inside the blade; the film gas holes adopt a funnel-shaped tapering hole structure, divided into a tapering section and a straight section, with the tapering section starting at 1 / 3 of the film gas hole length; the ratio of the diameter of the inlet to the outlet of the film gas hole is defined as the tapering ratio, ranging from 1.2 to 1.8; the angle formed by the axial direction of the film gas hole and the streamline direction is defined as the azimuth angle, ranging from 30° to 60°; the angle formed by the line connecting the centers of the outlets of the front and rear film gas holes and the streamline direction is defined as the deflection angle; the rear film gas holes are arranged in a direction ranging from -10° to -15° to avoid the deflection angle of the front film gas holes, with the angle pointing towards the blade root being positive.

Description

Technical Field

[0001] This invention belongs to the field of gas turbine rotor blade cooling technology, and in particular relates to a method for suppressing local failure of the gas film at the trailing edge of turbine blades. Background Technology

[0002] A gas turbine is a type of turbomachinery that exchanges energy with airflow. The turbine converts the extremely high enthalpy of the airflow into kinetic energy, which is then converted into mechanical energy by the turbine rotor blades. Under the impact of high-temperature, high-pressure gas, the mechanical properties of the metal materials used to manufacture the turbine rotor blades deteriorate sharply due to the increased temperature, making effective cooling of all parts of the blades crucial.

[0003] Turbine blades typically employ film cooling on their exterior. As the surface curvature and bending of turbine blades gradually increase along the outer radial direction, in order to form a uniform film on the blade surface, the azimuth angle of the film cooling holes is usually oriented towards the blade tip, and the film cooling holes on the suction surface are usually arranged in a longitudinal single-row or double-row configuration.

[0004] However, when the suction surface film gas holes are arranged in a longitudinal single-row or double-row configuration, the film gas coverage at the blade root is very poor. Related studies have shown that when the gas flows through the leading edge of the turbine rotor blade, horseshoe vortices are formed on both sides of the blade. The horseshoe vortex branches on the pressure surface merge with the low-momentum airflow near the endwall to form channel vortices. Due to the flow turning angle between the blades, a strong lateral pressure gradient is generated in the gas passage. Under the action of the lateral pressure gradient, the channel vortex moves downstream, leaving the leading edge of the pressure surface, approaching the suction surface of the adjacent blade, and finally attaching to the suction surface and flowing out of the passage.

[0005] However, the channel vortex near the suction surface rolls the air film of the rear half of the rotor blade root towards the middle of the blade, making it difficult for this area to be covered by the air film. This area eventually becomes the air film failure zone. The flow field characteristics of the air film failure zone make the air film cooling of the blade almost ineffective, resulting in extremely low cooling efficiency. This leads to poor blade structural strength in the air film failure zone. During engine operation, cracks will gradually appear in the blade structure in the air film failure zone, thereby shortening the service life of the turbine rotor blade.

[0006] Currently, turbine blade design does not consider how to suppress film failure, but solving the problem of film failure on the suction surface of the blade is crucial for the stable operation of high-performance gas turbines.

[0007] like Figure 1As shown, when the gas flows through the leading edge of the turbine rotor blade, horseshoe vortices are formed on both sides of the blade. The horseshoe vortices on the pressure surface of the blade merge with the low-energy fluid to form channel vortices. Due to the action of the lateral pressure gradient, the channel vortices will move downstream away from the pressure surface of the blade and move closer to the suction surface of the adjacent blade. Under the combined action of the angular vortex, the gas film in the rear half of the blade root is eventually rolled towards the middle of the blade, making it difficult for this area to be covered by the gas film.

[0008] like Figure 2 As shown, due to the low airflow velocity and high pressure on the suction surface of turbine rotor blades, separation and peeling phenomena are not easily generated. Usually, only one exhaust film hole is needed near the leading edge of the blade to form a stable air film coverage area, covering the entire suction surface of the blade up to the trailing edge. However, the entrainment effect of the channel vortex on the air film gradually increases in the rear half of the blade root, causing an approximately right-angled triangular air film failure area in the blade root region, which seriously affects the blade strength. Summary of the Invention

[0009] To address the problems existing in the prior art, this invention provides a method for suppressing localized film gas failure at the trailing edge of turbine blades. The method involves adjusting the arrangement of film gas holes in the film gas failure triangle region caused by the channel vortex, minimizing the overall impact on the blade. By using a funnel-shaped tapering orifice structure with increased azimuth, a velocity component opposite to the channel vortex is added to the outflow from the film gas holes, effectively suppressing the deviation of the film gas coverage area in the blade and increasing the spanwise and flowwise film gas coverage area. The film gas holes are arranged radially non-equidistantly according to the film gas coverage deflection angle, avoiding the influence of the upstream film gas coverage area on the outflow from the downstream film gas hole. The upstream film gas coverage area is located in the middle of the downstream film gas hole, solving the problem of no film gas coverage in the middle of the original film gas holes.

[0010] To achieve the above objectives, the present invention adopts the following technical solution: a method for suppressing local failure of the air film at the trailing edge of a turbine blade, specifically: dividing the trailing half of the blade root into a triangular area of ​​air film failure caused by the entrainment effect of the channel vortex on the air film, and adding non-uniformly arrayed air film holes in the triangular area of ​​air film failure.

[0011] The length of the air film failure triangle is 80% of the axial chord length from the trailing edge of the blade, and the height of the air film failure triangle is 35% of the leaf height from the leaf root.

[0012] The lateral spacing of the non-uniform array of film vents is set according to the position of the longitudinal air chamber inside the blade.

[0013] The non-uniform array arrangement of air film holes adopts a funnel-shaped tapering hole structure, which is divided into a tapering section and a straight section. The tapering section starts at 1 / 3 of the length of the air film hole. The ratio of the diameter of the air inlet to the air outlet of the non-uniform array arrangement of air film holes is defined as the tapering ratio, and the tapering ratio ranges from 1.2 to 1.8.

[0014] The angle formed by the axial direction of the non-uniform array of air film holes and the streamline direction is defined as the azimuth angle α, and the range of the azimuth angle α is 30° to 60°.

[0015] The angle formed by the line connecting the outlet centers of the non-uniform array of air film holes in the front row and the non-uniform array of air film holes in the rear row with the streamline direction is defined as the deflection angle β. The non-uniform array of air film holes in the rear row is arranged in a direction ranging from -10° to -15° to avoid the deflection angle β of the non-uniform array of air film holes in the front row, with the angle pointing towards the blade root being positive.

[0016] The beneficial effects of this invention are:

[0017] This invention modifies the air film pore structure on the blade's suction surface within a minimal range by dividing the air film failure triangle, thus minimizing the impact on blade strength.

[0018] This invention improves local cooling performance by adding air film holes towards the endwall in the air film failure triangle area at the leaf root, effectively resisting the influence of channel vortices, suppressing air film failure, and thus improving local cooling performance.

[0019] The air film pores of this invention abandon the conventional horizontal or vertical arrangement of air film pores and adopt a non-uniform array arrangement to make the air film coverage more uniform and achieve a larger air film coverage area with a limited number of pores.

[0020] This invention utilizes a funnel-shaped tapering orifice structure with an increased azimuth angle in the air film failure triangle region, enabling different air film orifice flow rates to precisely correspond to different channel vortices, thereby effectively enhancing air film efficiency. Attached Figure Description

[0021] Figure 1 A schematic diagram of the vortex in the blade cascade passage;

[0022] Figure 2 A schematic diagram showing the longitudinal single-row arrangement of the film cooling holes on the suction surface of the turbine blade;

[0023] Figure 3 A schematic diagram of adding non-uniform array arrangement of suction surface air film holes to the air film failure triangle region;

[0024] Figure 4 A schematic diagram of an air film pore structure employing a funnel-shaped tapering orifice and increasing the azimuth angle to address the air film failure triangle.

[0025] In the figure, 1—channel vortex, 2—corner vortex, 3—single row of air film holes, 4—air film coverage area, 5—air film failure triangle area, 6—non-uniform array of air film holes, 7—air inlet, 8—air outlet. Detailed Implementation

[0026] The present invention will now be described in further detail with reference to the accompanying drawings and specific embodiments.

[0027] A method for suppressing localized film gas failure at the trailing edge of turbine blades specifically involves: defining a film gas failure triangle 5 in the trailing half of the blade root, caused by the entrainment effect of the channel vortex 1 on the film gas; and adding a non-uniformly arrayed array of film gas holes 6 within the film gas failure triangle 5, such as... Figure 3 As shown.

[0028] The length of the air film failure triangle 5 is 80% of the axial chord length from the trailing edge of the blade, and the height of the air film failure triangle 5 is 35% of the leaf height from the leaf root.

[0029] The lateral spacing of the non-uniform array of air film holes 6 is set according to the position of the longitudinal air cavity inside the blade.

[0030] like Figure 4 As shown, the non-uniform array arrangement of air film holes 6 adopts a funnel-shaped tapering hole structure, which is divided into a tapering section and a straight section. The tapering section starts at 1 / 3 of the length of the air film hole. The ratio of the diameter of the air inlet 7 to the air outlet 8 of the non-uniform array arrangement of air film holes 6 is defined as the tapering ratio, and the tapering ratio ranges from 1.2 to 1.8.

[0031] The angle formed by the axial direction of the non-uniform array of air film holes 6 and the streamline direction is defined as the azimuth angle α, and the range of the azimuth angle α is 30° to 60°.

[0032] The angle formed by the line connecting the center of the air outlet 8 of the non-uniform array of air film holes 6 in the front row and the non-uniform array of air film holes 6 in the rear row with the streamline direction is defined as the deflection angle β. The non-uniform array of air film holes 6 in the rear row is arranged in a direction range of -10° to -15° to avoid the deflection angle β of the non-uniform array of air film holes 6 in the front row, with the angle pointing towards the blade root being positive.

[0033] by Figure 3 Taking the perspective as an example, the number of air film holes 6 arranged in a non-uniform array is five. The leftmost air film hole is defined as air film hole No. 1, the air film hole at the top of the middle side is defined as air film hole No. 2, the air film hole at the bottom of the middle side is defined as air film hole No. 3, the air film hole at the top of the rightmost side is defined as air film hole No. 4, and the air film hole at the bottom of the rightmost side is defined as air film hole No. 5.

[0034] Example 1:

[0035] In this embodiment, the turbine rotor blade chord length is 37 mm, the blade height (blade height) is 44 mm, the length of the film cooling failure triangle is 80% of the axial chord length from the blade trailing edge, and the height of the film cooling failure triangle is 35% of the blade height from the blade root. The straight section diameter of the non-uniformly arrayed film cooling holes 6 is 0.3 mm, the flow hole spacing is 7.0 mm, and the jet angle is 30°. The taper ratio of film cooling hole ① is 1.3, and the azimuth angle α is 50°. The taper ratio of film cooling hole ② is 1.2, the azimuth angle α is 45°, and the deflection angle β relative to film cooling hole ① is -20°. The taper ratio of film cooling hole ③ is 1.4, the azimuth angle α is 52°, and the deflection angle β relative to film cooling hole ① is 18°. The taper ratio of film cooling hole ④ is 1.2, the azimuth angle α is 41°, and the deflection angle β relative to film cooling hole ② is 7°. The taper ratio of air film pore No. 5 is 1.4, the azimuth angle α is 55°, and the deflection angle β relative to air film pore No. 3 is 8°.

[0036] Example 2:

[0037] In this embodiment, the turbine rotor blade chord length is 44 mm, the blade height (blade height) is 53 mm, the length of the film cooling failure triangle is 80% of the axial chord length from the blade trailing edge, and the height of the film cooling failure triangle is 35% of the blade height from the blade root. The straight section diameter of the non-uniformly arrayed film cooling holes 6 is 0.5 mm, the flow hole spacing is 8.0 mm, and the jet angle is 34°. The taper ratio of film cooling hole ① is 1.3, and the azimuth angle α is 52°. The taper ratio of film cooling hole ② is 1.2, the azimuth angle α is 42°, and the deflection angle β relative to film cooling hole ① is -23°. The taper ratio of film cooling hole ③ is 1.4, the azimuth angle α is 52°, and the deflection angle β relative to film cooling hole ① is 18°. The taper ratio of film cooling hole ④ is 1.2, the azimuth angle α is 41°, and the deflection angle β relative to film cooling hole ② is 2°. The taper ratio of air film pore No. 5 is 1.4, the azimuth angle α is 54°, and the deflection angle β relative to air film pore No. 3 is 3°.

[0038] Example 3:

[0039] In this embodiment, the turbine rotor blade chord length is 45 mm, the blade height (blade height) is 69 mm, the length of the film cooling failure triangle is 80% of the axial chord length from the blade trailing edge, and the height of the film cooling failure triangle is 35% of the blade height from the blade root. The straight section diameter of the non-uniformly arrayed film cooling holes 6 is 0.6 mm, the flow hole spacing is 8.5 mm, and the jet angle is 34°. The taper ratio of film cooling hole ① is 1.3, and the azimuth angle α is 45°. The taper ratio of film cooling hole ② is 1.2, the azimuth angle α is 41°, and the deflection angle β relative to film cooling hole ① is -20°. The taper ratio of film cooling hole ③ is 1.4, the azimuth angle α is 50°, and the deflection angle β relative to film cooling hole ① is 18°. The taper ratio of film cooling hole ④ is 1.2, the azimuth angle α is 38°, and the deflection angle β relative to film cooling hole ② is -7°. The taper ratio of air film pore No. 5 is 1.4, the azimuth angle α is 48°, and the deflection angle β relative to air film pore No. 3 is -8°.

[0040] The solutions described in the embodiments are not intended to limit the scope of patent protection of this invention. All equivalent implementations or modifications that do not depart from the scope of this invention are included in the patent scope of this case.

Claims

1. A method for suppressing localized failure of the gas film at the trailing edge of a turbine blade, characterized in that... Specifically, a triangular region of air film failure, caused by the entrainment effect of the channel vortex on the air film, is defined in the latter half of the blade root. Within this triangular region, a non-uniform array of air film holes is added. The non-uniform array of air film holes employs a funnel-shaped tapering hole structure, divided into a tapering section and a straight section. The tapering section begins at 1 / 3 of the air film hole length. The ratio of the diameter of the inlet to the outlet of the non-uniform array of air film holes is defined as the tapering ratio, with a range of 1.2–1.

8. The angle formed by the axial direction of the air film holes and the streamline direction is defined as the azimuth angle α, and the range of the azimuth angle α is 30° to 60°. The angle formed by the line connecting the outlet centers of the front row of non-uniform array air film holes and the rear row of non-uniform array air film holes and the streamline direction is defined as the deflection angle β. The rear row of non-uniform array air film holes are arranged to avoid the deflection angle β of the front row of non-uniform array air film holes in the range of -10° to -15°, where the angle pointing towards the blade root is positive. When the number of air film holes arranged in a non-uniform array is five, the leftmost air film hole is defined as air film hole No. 1, the air film hole at the top of the middle side is defined as air film hole No. 2, the air film hole at the bottom of the middle side is defined as air film hole No. 3, the air film hole at the top of the rightmost side is defined as air film hole No. 4, and the air film hole at the bottom of the rightmost side is defined as air film hole No.

5. Option 1: ① The taper ratio of air film aperture ① is 1.3, and the azimuth angle α is 50°; ② The taper ratio of air film aperture ② is 1.2, the azimuth angle α is 45°, and the deflection angle β relative to air film aperture ① is -20°; ③ The taper ratio of air film aperture ③ is 1.4, the azimuth angle α is 52°, and the deflection angle β relative to air film aperture ① is 18°; ④ The taper ratio of air film aperture ④ is 1.2, the azimuth angle α is 41°, and the deflection angle β relative to air film aperture ② is 7°; ⑤ The taper ratio of air film aperture ⑤ is 1.4, the azimuth angle α is 55°, and the deflection angle β relative to air film aperture ③ is 8°. Option 2: ① The taper ratio of air film aperture ① is 1.3, and the azimuth angle α is 52°; ② The taper ratio of air film aperture ② is 1.2, the azimuth angle α is 42°, and the deflection angle β relative to air film aperture ① is -23°; ③ The taper ratio of air film aperture ③ is 1.4, the azimuth angle α is 52°, and the deflection angle β relative to air film aperture ① is 18°; ④ The taper ratio of air film aperture ④ is 1.2, the azimuth angle α is 41°, and the deflection angle β relative to air film aperture ② is 2°; ⑤ The taper ratio of air film aperture ⑤ is 1.4, the azimuth angle α is 54°, and the deflection angle β relative to air film aperture ③ is 3°. Option 3: ① The taper ratio of air film aperture ① is 1.3, and the azimuth angle α is 45°; ② The taper ratio of air film aperture ② is 1.2, the azimuth angle α is 41°, and the deflection angle β relative to air film aperture ① is -20°; ③ The taper ratio of air film aperture ③ is 1.4, the azimuth angle α is 50°, and the deflection angle β relative to air film aperture ① is 18°; ④ The taper ratio of air film aperture ④ is 1.2, the azimuth angle α is 38°, and the deflection angle β relative to air film aperture ② is -7°; ⑤ The taper ratio of air film aperture ⑤ is 1.4, the azimuth angle α is 48°, and the deflection angle β relative to air film aperture ③ is -8°.

2. The method for suppressing localized film failure at the trailing edge of a turbine blade according to claim 1, characterized in that: The length of the air film failure triangle is 80% of the axial chord length from the trailing edge of the blade, and the height of the air film failure triangle is 35% of the leaf height from the leaf root.

3. The method for suppressing localized film failure at the trailing edge of a turbine blade according to claim 1, characterized in that: The lateral spacing of the non-uniform array of film vents is set according to the position of the longitudinal air chamber inside the blade.

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