A gas turbine, a gas turbine blade and a lobe-shaped film cooling hole structure thereof

Abstract

The application discloses a gas turbine, a gas turbine blade and a leaf-shaped film cooling hole structure of the gas turbine blade. The existing traditional cylindrical film cooling hole is reshaped into a leaf-shaped structure including a left half circular arc, a right upper half elliptical arc, a right middle half elliptical arc and a right lower half elliptical arc, so that the cooling gas can be divided into three paths when being sprayed out of the hole, and is sprayed out along the passage secondary flow same direction, the film hole center line direction and the passage secondary flow opposite direction respectively, thereby the blowing-off effect of the passage secondary flow on the cooling film is inhibited, the wall-attaching ability and the along-flow development ability of the cooling film are enhanced, the covering uniformity of the cooling film on the base surface is improved, and the cooling efficiency and the service life of the blade are improved. The leaf-shaped film cooling hole structure has the characteristics of strong functionality and wide applicability, and can be widely applied to the cooling technical field of the gas turbine blade of an aero-engine or a ground gas turbine.

Description

Technical Field

[0001] This invention belongs to the field of gas turbine blade cooling technology for aero-engines or ground gas turbines, and relates to a film cooling hole structure, specifically a blade-shaped film cooling hole structure suitable for gas turbine blades. Cooling gas is injected to the outside of the blade through the film cooling holes on the blade surface to form a cooling film covering the blade surface, thereby isolating the blade from the direct impact of the high-temperature main gas flow and realizing the cooling protection of the blade. Background Technology

[0002] Advanced civilian high-bypass turbofan engine turbine inlet temperatures exceed 2000K, far exceeding the maximum temperature that turbine blade materials can withstand. Film cooling (FSM) technology is an effective method that uses low-temperature cooling gas to form a protective layer on the turbine blade surface, thereby isolating the blades from thermal damage caused by the high-temperature main combustion flow. FSM can protect the blades and ensure their normal operation at high temperatures, preventing creep or damage due to high temperatures. The efficiency of FSM is not only affected by flow parameters such as air-to-air ratio, density ratio, and mainstream Reynolds number, but also closely related to geometric parameters.

[0003] Currently, researchers have conducted extensive studies on the shapes of film cooling orifices, including composite angle film cooling orifices and irregularly shaped orifices. Results show that optimizing and improving the geometry and related parameters of film cooling orifices can effectively enhance the film cooling effect compared to traditional cylindrical film cooling orifices. This is mainly because: during the mixing of the cryogenic cooling jet from a traditional cylindrical film cooling orifice with the high-temperature main combustion flow, the temperature and velocity difference between the jet and the main flow creates a kidney-shaped vortex near the wall, forcing the cryogenic cooling jet to detach from the wall while the high-temperature main combustion flow re-attaches to the wall. This results in a reduction in the film cooling coverage area and a decrease in the film cooling effect. By optimizing and improving the geometry and related parameters of the film cooling orifice, the exit area of ​​the cryogenic cooling jet can be increased to a certain extent, while the jet exit velocity can be reduced, thereby weakening the influence of the kidney-shaped vortex and ultimately improving the film cooling effect.

[0004] However, the existing studies mentioned above have focused more on the kidney-shaped vortex in the process of improving the mixing of the jet with the mainstream, and paid less attention to the impact of the secondary flow in the turbine blade passage. The secondary flow in the turbine blade passage refers to the vortex flow perpendicular to the mainstream direction generated in the turbine blade passage due to the boundary layer separation at the leading edge of the turbine blade and the lateral pressure gradient in the passage. Studies have shown that the secondary flow in the passage has an important impact on the film cooling technology of turbine blades, mainly in the following aspects: (1) The secondary flow in the passage will change the pressure distribution and velocity distribution at the outlet of the film cooling hole, thereby affecting the direction and angle of the film outflow. Generally speaking, secondary flow in the channel will cause the film gas outflow to deviate towards the suction side, which is beneficial to increase the film gas coverage area, but may also lead to an increase in mixing losses between the film gas outflow and the mainstream; (2) Secondary flow in the channel will also form a kidney-shaped vortex near the wall, forcing the low-temperature cooling jet to detach from the wall, and the high-temperature main combustion gas flow to reattach to the wall, resulting in a decrease in the film gas adhesion capacity; (3) Secondary flow in the channel will affect the film gas cooling effect, firstly affecting the distribution and coverage of the cooling gas on the blade surface, and secondly affecting the mixing process between the cooling gas and the main combustion gas flow; (4) Secondary flow in the channel will cause uneven temperature distribution on the blade surface, especially at the leading and trailing edges of the blade, increasing the thermal stress and thermal fatigue on the blade surface, and reducing the blade's life and reliability. In general, secondary flow in the channel with different intensities and directions can change the direction of film gas outflow and adhesion capacity, reduce the cooling effect, and is not conducive to the uniform temperature distribution on the blade surface.

[0005] In summary, film cooling technology is a key technology for high-temperature protection of advanced aero-engine turbine blades. Compared with other blade cooling technologies, it has advantages such as saving cooling gas, increasing blade life, and strong adaptability. Although existing film cooling technologies have made significant progress in structural design, there are still technical problems that need to be solved, such as how to improve the distribution of the kidney vortex to enhance the film adhesion to the wall, resist the negative effects of secondary flow in the channel, promote the development of the film along the flow direction, improve the uniformity of film coverage, and improve the cooling effect. It is necessary to comprehensively consider the influence of various flow physics phenomena and parameters, such as the flow inside the film orifice, the jet at the outlet of the film orifice, the mixing between the film outflow and the mainstream, the secondary flow in the turbine blade channel, and the geometry and arrangement of the film orifice. Summary of the Invention

[0006] (I) Purpose of the Invention

[0007] To address the aforementioned deficiencies and shortcomings of existing technologies, and to resolve the technical problems existing in using existing cylindrical film cooling holes, such as the kidney-shaped vortex lifting the cooled air and the secondary flow blowing away the cooled air in the channel leading to deterioration of the cooling effect, this invention aims to propose a gas turbine, gas turbine blades, and their leaf-shaped film cooling hole structure. By changing the film cooling hole structure from the traditional cylindrical shape to a leaf shape, the film cooling outflow is divided into three paths: one path adjusts the kidney-shaped vortex structure in the same direction as the secondary flow; the second path ensures the basic film cooling outflow along the flow direction; and the third path is used to resist the blowing away effect of the secondary flow in the channel. By adopting this leaf-shaped film cooling hole structure, the velocity and direction of the jet are changed at the outlet of the film cooling hole, which can improve the distribution of the kidney-shaped vortex and enhance the film cooling wall adhesion ability, while also resisting the negative effects of the secondary flow in the channel. Ultimately, it achieves the purpose of enhancing the development of the film cooling along the flow direction and improving the wall adhesion ability. It has the advantages of strong functionality and wide applicability, and has broad application prospects.

[0008] (II) Technical Solution

[0009] To achieve the objective of this invention, the present invention adopts the following technical solution:

[0010] The first objective of this invention is to provide a blade-shaped film cooling hole structure suitable for cooling gas turbine blades. The gas turbine blades are uniformly distributed within the thermal environment formed by the high-temperature main gas flow in an aero-engine or ground-based gas turbine, forming turbine blade channels. The gas turbine blades are hollow blade structures with a cryogenic cooling chamber, which is connected to an external low-temperature, high-pressure cooling gas source. The invention is characterized by...

[0011] The leaf-shaped film cooling holes are arranged in an array on the base of the gas turbine blade in two directions: the spanwise direction of the blade extension and the flow direction of the high-temperature main gas flow. The inlet end of each leaf-shaped film cooling hole is connected to the cryogenic cooling chamber of the gas turbine blade, and the outlet end extends to the base surface of the gas turbine blade in the high-temperature main gas flow environment. The leaf-shaped film cooling holes are used to spray the low-temperature high-pressure cooling gas in the cryogenic cooling chamber of the gas turbine blade into the thermal environment formed by the high-temperature main gas flow in the form of a jet, and form a cooling film covering the base surface of the gas turbine blade.

[0012] Each of the blade-shaped film cooling holes is machined and installed on the gas turbine blade with the hole centerline inclined relative to the base surface, and the angle θ between the hole centerline and the base surface is an acute angle. After each blade-shaped film cooling hole is projected onto the base surface, the length extension direction from the inlet end to the outlet end is basically consistent with the flow direction of the high-temperature main gas flow.

[0013] Each of the leaf-shaped film cooling holes has a leaf-shaped cross-section parallel to the substrate surface. The leaf-shaped hole structure is substantially symmetrical about its span relative to the projection of the hole's centerline onto the substrate surface. In the flow direction, it includes a left semicircular arc located upstream and a right upper branch semi-elliptical arc, a right middle branch semi-elliptical arc, and a right lower branch semi-elliptical arc located downstream and adjacent to the left semicircular arc. The upper branch semi-elliptical arc, the lower branch semi-elliptical arc, and the middle branch semi-elliptical arc are smoothly transitioned by circular arcs.

[0014] The diameter of the left semicircular arc is D, and the diameter D remains unchanged along the length extension direction of the leaf-shaped film cooling hole. The extension direction of the major axis of the upper right semi-elliptical arc is basically consistent with the flow direction of the secondary flow formed in the turbine blade passage. The extension direction of the major axis of the middle right semi-elliptical arc is basically consistent with the projection of the center line of the leaf-shaped film cooling hole onto the substrate surface. The extension direction of the major axis of the lower right semi-elliptical arc is basically consistent with the opposite direction of the flow direction of the secondary flow formed in the turbine blade passage.

[0015] The present invention relates to a leaf-shaped film cooling hole structure suitable for cooling gas turbine blades. Its working principle is as follows: the film cooling hole structure is modified from the existing traditional cylindrical structure into a leaf-shaped structure. During the process of low-temperature, high-pressure cooling gas being ejected through the leaf-shaped film cooling hole to form a low-temperature cooling jet, the outflow from the film cooling hole is divided into three paths. One path is ejected along the secondary flow of the channel via the upper right branch semi-elliptical arc to adjust the kidney-shaped vortex structure. The second path is ejected along the centerline of the film hole via the middle right branch semi-elliptical arc to ensure basic film outflow. The third path is ejected in the opposite direction along the secondary flow of the channel via the lower right branch semi-elliptical arc to resist the blowing effect of the secondary flow. The combined action of these three airflow paths ultimately enhances the development of the cooling film along the flow direction and improves its adhesion to the wall, fundamentally suppressing the blowing effect of the secondary flow of the channel on the cooling film outflow. This structure has the advantages of strong functionality and wide applicability.

[0016] Preferably, the angle θ between the center line of the leaf-shaped air film cooling hole and the substrate surface is an acute angle, and θ is between 20° and 60°.

[0017] Preferably, in the leaf-shaped air film cooling hole, the diameter D of the left semicircular arc remains unchanged along the length of the hole and is maintained between 0.8 and 4 mm.

[0018] Preferably, in the leaf-shaped air film cooling hole, the upper branch semi-elliptical arc and the middle branch semi-elliptical arc are smoothly transitioned by a circular arc with a radius of R1, where R1 is between 0.1D and 0.3D; the middle branch semi-elliptical arc and the lower branch semi-elliptical arc are smoothly transitioned by a circular arc with a radius of R2, where R2 is between 0.1D and 0.3D.

[0019] Preferably, in the leaf-shaped air film cooling hole, the major axis radius a1 of the upper right branch semi-elliptical arc is between 0.5D and 1D, and the minor axis radius b1 is about 0.2D; the major axis radius a2 of the middle right branch semi-elliptical arc is between 0.5D and 2D, and the minor axis radius b2 is between 0.25D and 0.5D; and the major axis radius a3 of the lower right branch semi-elliptical arc is between 0.5D and 1D, and the minor axis radius b3 is about 0.2D.

[0020] Furthermore, in the leaf-shaped air film cooling hole, the angle between the major axis of the upper semi-elliptical arc and the projection of the center line of the leaf-shaped air film cooling hole onto the substrate surface is α1, where α1 is between 0° and 75°; the angle between the major axis of the lower semi-elliptical arc and the projection of the center line of the leaf-shaped air film cooling hole onto the substrate surface is α2, where α2 is between 0° and 75°.

[0021] Furthermore, in the leaf-shaped film cooling hole, the included angles α1 and α2 are adjusted according to the flow direction of the secondary flow in the turbine blade channel, thereby adjusting the film outflow direction of the upper right semi-elliptical arc and the lower right semi-elliptical arc; and the major axis radius a1 and minor axis radius b1 of the upper right semi-elliptical arc and the major axis radius a3 and minor axis radius b3 of the lower right semi-elliptical arc are adjusted according to the intensity of the secondary flow in the turbine blade channel, thereby adjusting the film outflow intensity of the upper right semi-elliptical arc and the lower right semi-elliptical arc, ultimately resisting the adverse effects of the secondary flow in the turbine blade channel.

[0022] Furthermore, in the leaf-shaped air film cooling hole, the intensity of the basic air film outflow is adjusted by adjusting the major axis radius a2 of the right middle branch semi-elliptical arc between 0.5D and 2D and the minor axis radius b2.

[0023] Preferably, the gas turbine blade has a plurality of leaf-shaped film cooling holes arranged on its base along the spanwise direction of the blade, and the spacing between two adjacent leaf-shaped film cooling holes in the spanwise direction is P, where P is between 3D and 6D.

[0024] The second objective of this invention is to provide a gas turbine blade, wherein the gas turbine blade is uniformly distributed in the thermal environment formed by the high-temperature main gas flow in an aero-engine or a ground gas turbine and forms a turbine blade channel, characterized in that the gas turbine blade is a turbine guide vane or a turbine moving blade, and the base of the gas turbine blade is provided with leaf-shaped gas film cooling holes provided in the first objective of the invention arranged in an array.

[0025] The third objective of this invention is to provide a gas turbine, characterized in that the gas turbine is provided with the gas turbine blades provided in the second objective of the invention.

[0026] (III) Technical Effects

[0027] Compared with the prior art, the gas turbine, gas turbine blades, and leaf-shaped film cooling hole structure suitable for cooling gas turbine blades of the present invention have the following beneficial and significant technical effects:

[0028] (1) The present invention adopts a leaf-shaped air film cooling hole structure. The leaf-shaped air film hole is based on the existing traditional cylindrical air film cooling hole. The upstream half-cylindrical hole is retained, and the downstream half-cylindrical hole structure is modified into a leaf-shaped structure with three semi-elliptical arcs. The basic structure is relatively simple, and the three semi-elliptical arcs have different functions. When the cooling gas is ejected from the hole, it is ejected in the same direction as the secondary flow of the channel, in the direction of the center line of the air film hole, and in the opposite direction of the secondary flow of the channel. After combination, the kidney-shaped vortex structure can be improved, the active control of the secondary flow of the channel can be realized, the blowing effect of the secondary flow of the channel can be reduced, and the development of the cooling air film along the flow direction can be enhanced, thereby achieving the purpose of improving the cooling air film coverage and improving the cooling effect.

[0029] (2) The leaf-shaped film cooling hole structure of the present invention is suitable for cooling gas turbine blades. The leaf-shaped part of the leaf-shaped film cooling hole includes three elliptical arcs. The size and direction of the elliptical arcs determine the size and direction of the film outflow. The direction of the film outflow of the upper right semi-elliptical arc and the lower right semi-elliptical arc can be adjusted by flexibly adjusting the included angles α1 and α2 according to the direction of the secondary flow in the channel. The intensity of the film outflow of the upper right semi-elliptical arc and the lower right semi-elliptical arc can be adjusted by flexibly adjusting the size of the semi-major axis a1 and semi-minor axis b1, semi-major axis a3 and semi-minor axis b3 according to the intensity of the secondary flow in the channel. This achieves the optimized design of the direction and intensity of the cooling film outflow, and ultimately resists the adverse effects of the secondary flow in the channel. It suppresses the blowing effect of the secondary flow in the channel on the film outflow from the root, which is conducive to the development of the cooling film along the flow direction and improves the uniformity of the cooling film coverage on the surface of the machine body. Attached Figure Description

[0030] Figure 1The figure shows a top view (xz plane) of the blade-shaped film cooling hole structure applicable to gas turbine blade cooling of the present invention. The dashed line part in the figure is the internal structure of the hole.

[0031] Figure 2 This is an enlarged view of the leaf-shaped air film cooling hole of the present invention;

[0032] Figure 3 This is a cross-sectional view of the leaf-shaped air film cooling hole of the present invention.

[0033] Explanation of reference numerals in the attached figures:

[0034] 1-Base of gas turbine blade; 2-Leaf-shaped film cooling hole; 3-Left semicircular arc of leaf-shaped film cooling hole; 4-Right upper elliptical arc of leaf-shaped film cooling hole; 5-Right middle elliptical arc of leaf-shaped film cooling hole; 6-Right lower elliptical arc of leaf-shaped film cooling hole; 7-Inlet end of leaf-shaped film cooling hole; 8-Outlet end of leaf-shaped film cooling hole; 9-High-temperature main gas flow; 10-Low-temperature cooling jet; 11-Flow direction of high-temperature main gas flow; 12-Spanning direction of gas turbine blade; 13-Secondary flow in the channel. The x-axis is consistent with the spanning direction of the blade, the z-axis is consistent with the flow direction of the high-temperature main gas flow, and the y-axis is consistent with the thickness direction of the base. Detailed Implementation

[0035] To better understand the present invention, the following embodiments further illustrate its content. Throughout the accompanying drawings, the same or similar reference numerals denote the same or similar elements or elements having the same or similar functions. The described embodiments are some, but not all, of the embodiments of the present invention. The embodiments described below with reference to the accompanying drawings are exemplary and intended to explain the present invention, and should not be construed as limiting the present invention. Based on the embodiments of the present invention, all other embodiments obtained by those skilled in the art without creative effort are within the scope of protection of the present invention. The structure and technical solutions of the present invention will be further described in detail below with reference to the accompanying drawings, providing one embodiment of the present invention.

[0036] like Figures 1-3 As shown, the present invention provides a leaf-shaped film cooling hole structure suitable for cooling gas turbine blades. The gas turbine blades are uniformly distributed in the thermal environment formed by the high-temperature main gas flow in the aero-engine or ground gas turbine and form a turbine blade channel. The gas turbine blades are hollow blade structures with a low-temperature cold air cavity, which is connected to an external low-temperature high-pressure cold air source.

[0037] Leaf-shaped film cooling holes 2 are arranged in an array on the base 1 of the gas turbine blade in two directions: the spanwise direction 12 of the blade extension and the flow direction 11 of the high-temperature main gas flow 9. The inlet end 7 of each leaf-shaped film cooling hole 2 is connected to the cryogenic cold air chamber of the gas turbine blade, and the outlet end 8 extends to the surface of the base 1 of the gas turbine blade in the high-temperature main gas flow environment. The leaf-shaped film cooling holes 2 are used to spray the low-temperature high-pressure cold air in the cryogenic cold air chamber of the gas turbine blade into the thermal environment formed by the high-temperature main gas flow 9 in the form of a jet 10, and form a cooling film covering the surface of the base 1 of the gas turbine blade. Each leaf-shaped film cooling hole 2 is machined and installed on the gas turbine blade with the hole centerline inclined relative to the surface of the substrate 1, and the angle θ between the hole centerline and the surface of the substrate 1 is an acute angle, preferably between 20° and 60°. After each leaf-shaped film cooling hole 2 is projected onto the surface of the substrate 1, the length extension direction between its inlet end and outlet end is basically consistent with the flow direction 11 of the high-temperature main gas flow 9.

[0038] Each leaf-shaped air film cooling hole 2 has a leaf-shaped hole structure in its cross-section parallel to the substrate surface. The leaf-shaped hole structure is basically symmetrical about the projection of the hole centerline onto the substrate surface in the spanwise direction 12. In the flow direction, it includes a left semicircular arc 3 located upstream and a right upper branch semi-elliptical arc 4, a right middle branch semi-elliptical arc 5, and a right lower branch semi-elliptical arc 6 located downstream and adjacent to the left semicircular arc 3. The upper branch semi-elliptical arc 4, the lower branch semi-elliptical arc 6 and the middle branch semi-elliptical arc 5 are smoothly transitioned by arcs. The radius of the arc between the upper branch semi-elliptical arc 4 and the middle branch semi-elliptical arc 5 is R1, preferably between 0.1D and 0.3D. The radius of the arc between the middle branch semi-elliptical arc 5 and the lower branch semi-elliptical arc 6 is preferably between 0.1D and 0.3D. Furthermore, the diameter of the left semicircular arc 3 is D, which remains unchanged along the length extension direction of the blade-shaped film cooling hole, and preferably is kept between 0.8 and 4 mm. The extension direction of the major axis of the upper right semi-elliptical arc 4 is basically consistent with the flow direction of the secondary flow formed in the turbine blade channel. The extension direction of the major axis of the middle right semi-elliptical arc 5 is basically consistent with the projection of the center line of the blade-shaped film cooling hole on the substrate surface. The extension direction of the major axis of the lower right semi-elliptical arc 6 is basically consistent with the opposite direction of the flow direction of the secondary flow formed in the turbine blade channel.

[0039] Furthermore, in the leaf-shaped film cooling hole 2, the major axis radius a1 of the upper right semi-elliptical arc 4 is between 0.5D and 1D, and the minor axis radius b1 is approximately 0.2D; the major axis radius a2 of the middle right semi-elliptical arc 5 is between 0.5D and 2D, and the minor axis radius b2 is between 0.25D and 0.5D; the major axis radius a3 of the lower right semi-elliptical arc 6 is between 0.5D and 1D, and the minor axis radius b3 is approximately 0.2D. The angle between the major axis of the upper semi-elliptical arc 4 and the projection of the center line of the leaf-shaped film cooling hole onto the substrate surface is α1, where α1 is between 0° and 75°; the angle between the major axis of the lower semi-elliptical arc 6 and the projection of the center line of the leaf-shaped film cooling hole onto the substrate surface is α2, where α2 is between 0° and 75°. Furthermore, in the blade-shaped film cooling hole 2, the included angles α1 and α2 are adjusted according to the flow direction of the secondary flow within the turbine blade passage, thereby adjusting the film outflow direction of the upper right semi-elliptical arc 4 and the lower right semi-elliptical arc 6. The major axis radius a1 and minor axis radius b1 of the upper right semi-elliptical arc 4 and the major axis radius a3 and minor axis radius b3 of the lower right semi-elliptical arc 6 are adjusted according to the intensity of the secondary flow within the turbine blade passage, thereby adjusting the film outflow intensity of the upper right semi-elliptical arc 4 and the lower right semi-elliptical arc 6, ultimately resisting the adverse effects of the secondary flow within the turbine blade passage. By adjusting the major axis radius a2 of the middle right semi-elliptical arc 5 to between 0.5D and 2D, and the minor axis radius b2, the intensity of the basic film outflow is adjusted.

[0040] The leaf-shaped film cooling hole structure of the present invention, applicable to the cooling of gas turbine blades, works on the following principle: the film cooling hole structure is modified from the existing traditional cylindrical structure into a leaf-shaped structure, so that during the process of low-temperature and high-pressure cooling gas being ejected through the leaf-shaped film cooling hole to form a low-temperature cooling jet, the outflow of the film cooling hole is divided into three paths. One path is ejected along the secondary flow of the channel in the same direction as the upper right branch semi-elliptical arc 4 to adjust the kidney-shaped vortex structure. The second path is ejected along the center line of the film hole through the middle right branch semi-elliptical arc 5 to ensure the basic film outflow. The third path is ejected in the opposite direction along the secondary flow of the channel through the lower right branch semi-elliptical arc 6 to resist the blowing effect of the secondary flow of the channel. The three airflows work together to ultimately enhance the development of the cooling film along the flow direction and improve its adhesion ability, thereby fundamentally suppressing the blowing effect of the secondary flow of the channel on the cooling film outflow.

[0041] Compared with existing technologies, this invention adopts a leaf-shaped air film cooling hole structure. The leaf-shaped air film cooling hole is based on the existing traditional cylindrical air film cooling hole. The upstream semi-cylindrical hole is retained, and the downstream semi-cylindrical hole structure is modified into a leaf-shaped structure with three semi-elliptical arcs. The size and direction of the elliptical arcs determine the size and direction of the air film outflow. The outflow direction of the upper and lower right semi-elliptical arcs can be adjusted by flexibly adjusting the included angles α1 and α2 according to the direction of the secondary flow in the channel. The intensity of the outflow of the air film in the upper and lower right semi-elliptical arcs can be adjusted by flexibly adjusting the size of the semi-major axis a1 and semi-minor axis b1, semi-major axis a3 and semi-minor axis b3 according to the intensity of the secondary flow in the channel. This achieves optimized design of the direction and intensity of the cooling air film outflow, and ultimately resists the adverse effects of the secondary flow in the channel. It fundamentally suppresses the blowing effect of the secondary flow in the channel on the air film outflow, which is conducive to the development of the cooling air film along the flow direction and improves the uniformity of the cooling air film coverage on the machine surface.

[0042] The objectives of this invention have been fully and effectively achieved through the above embodiments. Those skilled in the art will understand that this invention includes, but is not limited to, the contents described in the accompanying drawings and the specific embodiments described above. Although the invention has been described with reference to what is currently considered the most practical and preferred embodiments, it should be understood that the invention is not limited to the disclosed embodiments, and any modifications that do not depart from the functional and structural principles of the invention will be included within the scope of the claims.

Claims

1. A blade-shaped film cooling hole structure suitable for cooling gas turbine blades, wherein the gas turbine blades are uniformly distributed in the thermal environment formed by the high-temperature main gas flow in an aero-engine or ground gas turbine and form turbine blade channels, wherein the gas turbine blades are hollow blade structures with a cryogenic cooling chamber, the cryogenic cooling chamber being connected to an external cryogenic high-pressure cooling gas source, characterized in that, The leaf-shaped film cooling holes are arranged in an array on the substrate of the gas turbine blade in two directions: the spanwise direction of the blade extension and the flow direction of the high-temperature main gas flow. The inlet end of each leaf-shaped film cooling hole is connected to the cryogenic cooling chamber of the gas turbine blade, and the outlet end extends to the substrate surface of the gas turbine blade in the high-temperature main gas flow environment. The leaf-shaped film cooling holes are used to spray the low-temperature high-pressure cooling gas in the cryogenic cooling chamber of the gas turbine blade into the thermal environment formed by the high-temperature main gas flow in the form of a jet, and form a cooling film covering the substrate surface of the gas turbine blade. Each of the aforementioned leaf-shaped film cooling holes is machined on the gas turbine blade with its centerline inclined relative to the substrate surface, and the angle between the centerline of the hole and the substrate surface is... θ The angle is acute, and after projecting each leaf-shaped air film cooling hole onto the substrate surface, the length extension direction between its inlet end and outlet end is consistent with the flow direction of the high-temperature main combustion gas flow. Each leaf-shaped film cooling hole has a leaf-shaped cross-section parallel to the substrate surface. The leaf-shaped hole structure is symmetrical in the spanwise direction with respect to the projection of the hole's centerline onto the substrate surface. In the flow direction, it includes a left semicircular arc located upstream and a right upper branch semi-elliptical arc, a right middle branch semi-elliptical arc, and a right lower branch semi-elliptical arc located downstream and adjacent to the left semicircular arc. The upper branch semi-elliptical arc, the lower branch semi-elliptical arc, and the middle branch semi-elliptical arc are smoothly transitioned by circular arcs. The diameter of the left semicircular arc is D, and D remains unchanged along the length of the leaf-shaped film cooling hole. The major axis of the upper right semi-elliptical arc extends in the same direction as the secondary flow formed in the turbine blade passage. The major axis of the middle right semi-elliptical arc extends in the same direction as the projection of the center line of the leaf-shaped film cooling hole onto the substrate surface. The major axis of the lower right semi-elliptical arc extends in the opposite direction to the secondary flow formed in the turbine blade passage. The angle between the major axis of the upper semi-elliptical arc and the projection of the center line of the blade-shaped film cooling hole onto the base surface is α1, where α1 is between 0° and 75°; the angle between the major axis of the lower semi-elliptical arc and the projection of the center line of the blade-shaped film cooling hole onto the base surface is α2, where α2 is between 0° and 75°; the angles α1 and α2 are determined according to the flow direction of the secondary flow within the turbine blade passage to adjust the film outflow direction of the upper right semi-elliptical arc and the lower right semi-elliptical arc; the upper right semi-elliptical arc has a major axis radius a1 and a minor axis radius b1, and the lower right semi-elliptical arc... The elliptical arc has a major axis radius a3 and a minor axis radius b3. The major axis radius a1, minor axis radius b1, major axis radius a3, and minor axis radius b3 are determined according to the intensity of the secondary flow in the turbine blade passage to adjust the film outflow intensity of the upper right semi-elliptical arc and the lower right semi-elliptical arc. The cooling gas ejected through the upper right semi-elliptical arc is injected in the same direction as the secondary flow in the passage, the cooling gas ejected through the middle right semi-elliptical arc is injected in the direction of the center line of the film orifice, and the cooling gas ejected through the lower right semi-elliptical arc is injected in the opposite direction as the secondary flow in the passage, so as to jointly resist the blowing effect of the secondary flow in the turbine blade passage on the cooling film.

2. The blade-shaped film cooling hole structure for gas turbine blade cooling according to claim 1, characterized in that, The angle between the centerline of the leaf-shaped air film cooling hole and the substrate surface θ An acute angle, θ Between 20° and 60°.

3. The blade-shaped film cooling hole structure for gas turbine blade cooling according to claim 1, characterized in that, In the leaf-shaped air film cooling hole, the diameter D of the left semicircular arc remains unchanged along the length of the hole and is maintained between 0.8 and 4 mm.

4. The blade-shaped film cooling hole structure for gas turbine blade cooling according to claim 1, characterized in that, In the leaf-shaped air film cooling hole, the upper branch semi-elliptical arc and the middle branch semi-elliptical arc are smoothly transitioned by a circular arc with a radius of R1, which is between 0.1D and 0.3D; the middle branch semi-elliptical arc and the lower branch semi-elliptical arc are smoothly transitioned by a circular arc with a radius of R2, which is between 0.1D and 0.3D.

5. The blade-shaped film cooling hole structure for gas turbine blade cooling according to claim 1, characterized in that, In the leaf-shaped air film cooling hole, the major axis radius a1 of the upper right semi-elliptical arc is between 0.5D and 1D, and the minor axis radius b1 is about 0.2D; the major axis radius a2 of the middle right semi-elliptical arc is between 0.5D and 2D, and the minor axis radius b2 is between 0.25D and 0.5D; the major axis radius a3 of the lower right semi-elliptical arc is between 0.5D and 1D, and the minor axis radius b3 is about 0.2D.

6. The blade-shaped film cooling hole structure for gas turbine blade cooling according to claim 1, characterized in that, The gas turbine blade has multiple leaf-shaped film cooling holes arranged along the span of the blade on its base. The spacing between two adjacent leaf-shaped film cooling holes in the span is P, where P is between 3D and 6D.

7. A gas turbine blade, wherein the gas turbine blades are uniformly distributed in the thermal environment formed by the high-temperature main gas flow in an aero-engine or a ground-based gas turbine and form a turbine blade passage, characterized in that, The gas turbine blade is a turbine guide vane or a turbine moving blade, and the base of the gas turbine blade is arranged in an array with the leaf-shaped gas film cooling hole structure as described in any one of claims 1 to 6, which is suitable for cooling gas turbine blades.

8. A gas turbine, characterized in that, The gas turbine blades as described in claim 7 are provided.

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