A turbine blade film cooling structure for an aeroengine

Abstract

The application discloses a turbine blade film cooling structure of an aero-engine, wherein a film cooling hole in the turbine blade film cooling structure is provided with a double-C-shaped expansion film hole structure. The structure aims at the technical problem that a kidney vortex is easily formed near an outlet of a film hole on a blade surface in a prior art cylindrical film hole, thereby causing the cold gas to be lifted, the coverage area of the film to be reduced, and the cooling effect to be deteriorated. The double-C-shaped expansion film hole is adopted to construct an anti-kidney vortex near the outlet of the film hole to inhibit the formation of the kidney vortex and the action of the kidney vortex, so that the purpose of enhancing the film wall adhesion is achieved. Meanwhile, the aerodynamic loss is not obviously increased, and the double-C-shaped expansion film hole has the advantages of simple structure and wide applicability. The processing difficulty of the double-C-shaped expansion film hole is not much different from that of the prior art cylindrical film hole. The film cooling effect can be greatly improved, and the film cooling effect can be greatly improved. The double-C-shaped expansion film hole has a wide application prospect.

Description

Technical Field

[0001] This invention belongs to the field of aero-engine turbine blade cooling technology, and relates to an aero-engine turbine blade film cooling structure. Specifically, it relates to a turbine blade film cooling structure with double C-shaped expansion film cooling holes. The double C-shaped expansion film cooling holes arranged front and rear in this invention are beneficial for the rapid formation of anti-kidney vortices on the turbine blade surface, suppressing kidney vortices and enhancing the adhesion ability of the film. At the same time, there is no significant increase in aerodynamic losses. Moreover, the overall structure of the double C-shaped expansion film cooling holes is relatively simple, and the processing difficulty is not much different from that of existing cylindrical film cooling holes. The film cooling effect can be greatly improved, and it has broad application prospects. Background Technology

[0002] Turbine blades are the components with the highest thermal and mechanical loads in aero-turbine engines. Ablation, cracking, spalling, and fracture are common turbine blade failures, and these failures are almost all closely related to the thermal conditions of the turbine blades. The inlet temperature of turbines in advanced civil high-bypass aero-engines has exceeded 2000K, which is far higher than the highest temperature that current turbine blade materials can withstand. To ensure the safe and reliable operation of turbine blades, the heat transfer design currently mainly adopts a combination of methods such as disturbance convection cooling, impact cooling, film cooling, sweating cooling, plate cooling, and thermal barrier coatings to achieve turbine blade cooling. Film cooling is one of the main cooling methods for high-temperature components in gas turbines due to its high cooling efficiency and relatively flexible arrangement. Film cooling technology can protect the blades from normal operation in the high-temperature main gas flow environment and prevent the blades from creeping or being damaged due to high temperature. In film cooling (FSL) technology, a cryogenic cooling jet enters the high-temperature main combustion gas flow region through holes or slots on the blade surface along a specific injection direction. Due to the pressure of the high-temperature main combustion gas flow and the friction of the wall, the cryogenic cooling jet covers the blade surface, isolating the high-temperature main combustion gas flow from the blade, weakening the heat exchange process between the high-temperature main combustion gas flow and the blade, thereby reducing the blade temperature and preventing corrosion of the blade by impurities in the high-temperature combustion gas. The efficiency of FSL is not only affected by flow parameters such as the blowing ratio and density ratio of the cryogenic cooling jet, and the Reynolds number of the main flow of the high-temperature combustion gas, but also closely related to the geometry and parameters of the FSL orifice.

[0003] Currently, researchers have conducted extensive studies on the geometry optimization of film cooling orifices, including composite angle film cooling orifices and irregularly shaped orifices. Results show that improving the combined shape of the film cooling orifice can effectively enhance the film cooling effect compared to the commonly used cylindrical film cooling orifice. This is mainly because, during the mixing process between the jet and the mainstream in a cylindrical film cooling orifice, due to the temperature and velocity differences between the jet and the mainstream, a layer of material forms on the turbine blade wall to be cooled, such as… Figure 3The kidney-shaped vortex 10 shown forces the cryogenic cooling jet to detach from the turbine blade wall, allowing the high-temperature main combustion gas to re-attach to the blade wall. This reduces the film cooling area on the blade wall, thus decreasing the film cooling effect. Improving the geometry of the film cooling orifice and increasing its exit area to a certain extent can reduce the jet exit velocity, which helps to weaken the formation and influence of the kidney-shaped vortex, thereby improving the film cooling effect. However, irregularly shaped orifices and other film cooling orifice structures are relatively complex, difficult to manufacture, and costly. They also affect the mainstream aerodynamic performance of the turbine blade surface and increase aerodynamic losses. Summary of the Invention

[0004] (I) Purpose of the Invention

[0005] To address the aforementioned deficiencies and shortcomings of existing technologies, the present invention aims to propose a film cooling structure for aero-engine turbine blades. This structure features double-C-shaped expanding film cooling holes, which address the technical problem of existing cylindrical film cooling holes easily forming kidney-shaped vortices near the outlet on the blade surface. This leads to the lifting of cooled air, reduced film coverage area, and deterioration of cooling effect. By employing double-C-shaped expanding film cooling holes to construct anti-kidney-shaped vortices near the outlet, the formation and effect of kidney-shaped vortices are suppressed, thereby enhancing film adhesion. Simultaneously, aerodynamic losses are not significantly increased. Furthermore, the double-C-shaped expanding film cooling holes have the advantages of simple structure and wide applicability. Their manufacturing difficulty is not significantly different from existing cylindrical film cooling holes, and the film cooling effect can be greatly improved, demonstrating broad application prospects.

[0006] (II) Technical Solution

[0007] To achieve its inventive purpose and solve its technical problems, the technical solution adopted by this invention is as follows:

[0008] A film cooling structure for an aero-engine turbine blade, wherein the turbine blade is a hollow blade, and its internal cavity is filled with low-temperature cooling high-pressure gas introduced from the outside. Multiple pairs of double-C-shaped expansion film cooling holes communicating with the internal cavity are arranged on the blade wall substrate. The outside of the turbine blade is a high-temperature environment formed by the high-temperature main combustion gas flow. The film cooling holes are used to inject the high-pressure gas from the internal cavity of the blade into the high-temperature environment as a low-temperature cooling jet, forming a cooling film covering the outer wall of the turbine blade. The structure is characterized by...

[0009] Each of the aforementioned double C-shaped expansion film cooling hole pairs includes an upstream C-shaped expansion film cooling hole and a downstream C-shaped expansion film cooling hole. The upstream and downstream C-shaped expansion film cooling holes are arranged in a front-to-back manner along the mainstream direction of the high-temperature main combustion flow. The line connecting their axes is parallel to the mainstream direction of the high-temperature main combustion flow. The inlet ends of the upstream and downstream C-shaped expansion film cooling holes are both formed on the inner wall surface of the turbine blade wall substrate, and the outlet ends are both formed on the outer wall surface of the turbine blade wall substrate. The hole spacing e between the inlet ends of the upstream and downstream C-shaped expansion film cooling holes is equivalent to the hole spacing f between their outlet ends.

[0010] The cross-sections of the upstream C-shaped expansion film orifice and the downstream C-shaped expansion film orifice are both semi-elliptical arc-shaped slits, and the major axis of the semi-elliptical arc-shaped slit is basically perpendicular to the mainstream direction of the high-temperature main gas flow, and the minor axis is basically parallel to the mainstream direction of the high-temperature main gas flow.

[0011] The upstream and downstream C-shaped expansion film orifices each include an inner ring sidewall located radially inward and an outer ring sidewall located radially outward. At the outlet end, the major axis radius of the outer ring sidewall is a1 and the minor axis radius is b1, while the major axis radius of the inner ring sidewall is a2 and the minor axis radius is b2. At the inlet end, the major axis radius of the outer ring sidewall is a3 and the minor axis radius is b3, while the major axis radius of the inner ring sidewall is a4 and the minor axis radius is b4. Furthermore, a4 ≤ a2, b4 ≤ b2, a3 ≤ a1, and b3 ≤ b1.

[0012] The inner and outer ring sidewalls of the upstream and downstream C-shaped expanding air film holes are both expanding with an expansion angle α from their inlet to outlet ends.

[0013] Both the upstream and downstream C-shaped expansion film air holes are inclinedly arranged on the blade wall substrate, and the angle θ between their center lines and the tangent direction of the blade wall substrate is an acute angle.

[0014] In a preferred embodiment of the present invention, the pairs of double C-shaped expansion film cooling holes are arranged in an array on the blade wall substrate, and the hole spacing P between two adjacent pairs of double C-shaped expansion film cooling holes in the direction perpendicular to the main flow of the high-temperature main gas flow is between 3a1 and 6a1.

[0015] In a preferred embodiment of the present invention, the angle θ between the centerlines of the upstream C-shaped expansion film pore and the downstream C-shaped expansion film pore and the tangent direction of the blade wall substrate is in the range of 20 to 60°.

[0016] In a preferred embodiment of the present invention, the expansion angle α between the inner and outer ring sidewalls of the upstream and downstream C-shaped expansion film orifices, from their inlet to outlet ends, ranges from 0 to 6°.

[0017] In a preferred embodiment of the present invention, at the air outlet end, the major axis radius a1 of the outer ring sidewall is in the range of 4 to 10 mm, the minor axis radius b1 is in the range of 2 to 5 mm, and the major axis radius a2 of the inner ring sidewall is in the range of 3 to 7 mm, the minor axis radius b2 is in the range of 1 to 2 mm.

[0018] In a preferred embodiment of the present invention, at the intake end, the major axis radius a3 of the outer ring sidewall ranges from 1.6 to 4 mm, the minor axis radius b3 ranges from 0.4 to 2 mm, and the major axis radius a4 of the inner ring sidewall ranges from 1.2 to 3 mm, the minor axis radius b4 ranges from 0.4 to 1 mm.

[0019] In a preferred embodiment of the present invention, the numerical range of the hole spacing e between the inlet ends of the upstream C-shaped expansion air film hole and the downstream C-shaped expansion air film hole is between a1 and 2a1, and the numerical range of the hole spacing f between their outlet ends is between a1 and 2a1.

[0020] The working principle of the air film cooling structure for turbine blades of the aero-engine of the present invention is as follows:

[0021] To address the technical problem of existing turbine blade film cooling structures using cylindrical film cooling holes, which easily form kidney-shaped vortices near the outlet of the film cooling hole on the blade surface, leading to cold gas lifting, reduced film coverage area, and deterioration of film cooling effect, this invention proposes a turbine blade film cooling structure that replaces the existing cylindrical film cooling holes with double C-shaped expansion film cooling holes. On the one hand, the cross-section of the upstream and downstream C-shaped expansion film cooling holes in the double C-shaped expansion film cooling holes is a semi-elliptical arc-shaped slit, and the major axis of the semi-elliptical arc-shaped slit is basically perpendicular to the mainstream direction of the high-temperature main gas flow, while the minor axis is basically parallel to the mainstream direction of the high-temperature main gas flow. Since the size of the double C-shaped expansion film cooling holes in the direction perpendicular to the mainstream direction is larger than that in the direction parallel to the mainstream direction, the lateral coverage width and coverage area of ​​the film formed on the surface of the turbine blade to be cooled are increased, thereby improving the film cooling effect to a certain extent.

[0022] On the other hand, and more importantly, replacing the existing cylindrical film cooling orifices with double C-shaped expansion film cooling orifices results in a significant increase in pressure. Since both the upstream and downstream C-shaped expansion orifices in the double C-shaped orifices have an expansion angle α (0–6°) at their outlets, the flow velocity decreases and the pressure increases as the cryogenic cooling flow passes through the outlet with the expansion angle α. This creates a pair of vortices near the orifice outlets on the blade surface, opposite in direction to the kidney-shaped vortex rotation—the anti-kidney vortex. The anti-kidney vortex can counteract or partially counteract the effect of the kidney-shaped vortex because they have opposite directions and velocities. This reduces the longitudinal height and increases the lateral width of the kidney-shaped vortex, suppressing the adverse effects of the kidney-shaped vortex, such as gas flow instability, uneven pressure distribution in the film cooling system, and reduced film cooling efficiency. This also greatly increases the coverage area of ​​the film cooling system on the turbine blade surface to be cooled. Furthermore, the generation of the anti-kidney vortex helps reduce the variation range of the film cooling flow rate, thereby improving the uniformity and consistency of film cooling. Because the formation of the inverted kidney vortex can balance the momentum and mass transfer in the gas flow, reducing the range of gas flow variation, thereby reducing the inhomogeneity and fluctuation of the gas film. It should also be noted that, in this invention, by adjusting the hole spacing e at the inlet end and the hole spacing f at the outlet end of the upstream and downstream C-shaped expansion gas film orifices, the scale of the inverted kidney vortex can be easily adjusted, optimizing the gas flow distribution and velocity, thereby controlling the gas film coverage and cooling effect, and improving the efficiency, uniformity, and stability of gas film cooling.

[0023] Finally, it should be noted that the present invention uses double C-shaped expanding air film holes to construct anti-kidney-shaped vortices near the outlet of the air film holes to suppress the formation and effect of kidney-shaped vortices, thereby achieving the purpose of enhancing air film adhesion. At the same time, there is no significant increase in aerodynamic losses. Moreover, the double C-shaped expanding air film holes have the advantages of simple structure and wide applicability. Their processing difficulty is not much different from that of existing cylindrical air film holes. The film cooling effect can be greatly improved, and it has broad application prospects.

[0024] (III) Technical Effects

[0025] Compared with the prior art, the air film cooling structure for aero-engine turbine blades with double C-shaped expansion air film holes provided by the present invention has the following characteristics and significant technical effects:

[0026] (1) Unique structure: In the air film cooling structure of the turbine blade of the aero-engine of the present invention, the double C-shaped expansion air film hole includes a pair of expansion air film holes arranged in front and behind, without relying on any existing air film hole structure. The overall structure is relatively simple and the requirements for processing technology are not high.

[0027] (2) Good adjustability: In the air film cooling structure of the turbine blade of the aero-engine of the present invention, the structure of the kidney vortex and the anti-kidney vortex can be directly changed by adjusting the distance between the upstream hole and the downstream hole, so as to obtain good cooling effect on different applications.

[0028] (3) Wide range of applications: The air film cooling structure for turbine blades of the present invention can be used for hot end components of aero engines such as turbine guide vanes, moving blades, and combustion chambers to achieve efficient cooling for different needs.

[0029] (4) Significant improvement in cooling effect: The air film cooling structure for turbine blades of the aero-engine of the present invention suppresses the negative effects of the kidney vortex from the source, and the cooling effect is significantly improved. Attached Figure Description

[0030] Figure 1 This is a top view of the air film cooling structure for aero-engine turbine blades according to the present invention.

[0031] Figure 2 This is a cross-sectional view of the double C-shaped expanding air film pore pair in this invention.

[0032] Figure 3 This is a schematic diagram of an existing cylindrical air-film pore kidney-shaped vortex structure.

[0033] Figure 4 This is a schematic diagram of the double C-shaped expanding air membrane pore kidney-shaped vortex and the anti-kidney-shaped vortex structure.

[0034] Explanation of reference numerals in the attached figures:

[0035] 1- Turbine blade wall substrate; 2- Upstream C-shaped expanding film gas hole; 3- Downstream C-shaped expanding film gas hole; 4- Inner ring sidewall; 5- Outer ring sidewall; 6- Inlet end; 7- Outlet end; 8- Main stream; 9- Jet; 10- Kidney-shaped vortex formed by existing cylindrical film gas holes; 11- Kidney-shaped vortex formed by a pair of double C-shaped expanding film gas holes; 12- Anti-kidney-shaped vortex formed by a pair of double C-shaped expanding film gas holes; e- Spacing between the inlet ends of the upstream and downstream C-shaped expanding film gas holes; f- Spacing between the upstream and downstream C-shaped expanding film gas holes; a1- Major axis radius of the outlet end of the outer ring sidewall; b1- Minor axis radius of the outlet end of the outer ring sidewall; a2- Major axis radius of the outlet end of the inner ring sidewall; b2- Minor axis radius of the outlet end of the inner ring sidewall; a3- Major axis radius of the inlet end of the outer ring sidewall. Radius, b3 - minor axis radius of the inlet end of the outer ring sidewall, a4 - major axis radius of the inlet end of the inner ring sidewall, b4 - minor axis radius of the inlet end of the inner ring sidewall, θ - angle between the center lines of the upstream and downstream C-shaped expanding film air holes and the tangent direction of the blade wall matrix, α - expansion angle between the inner and outer ring sidewalls from their inlet to outlet ends, m1 - height dimension of the kidney-shaped vortex formed by the existing cylindrical film air holes, m2 - height dimension of the kidney-shaped vortex formed by the pair of double C-shaped expanding film air holes, m3 - height dimension of the anti-kidney-shaped vortex formed by the pair of double C-shaped expanding film air holes, n1 - width dimension of the kidney-shaped vortex formed by the existing cylindrical film air holes, n2 - width dimension of the kidney-shaped vortex formed by the pair of double C-shaped expanding film air holes, n3 - width dimension of the anti-kidney-shaped vortex formed by the pair of double C-shaped expanding film air holes. Detailed Implementation

[0036] To better understand the present invention, the following embodiments further illustrate the content of the invention, so that the advantages and features of the invention can be more easily understood by those skilled in the art. It should be noted that the following descriptions are merely preferred embodiments of the present invention, but the content of the invention is not limited to the following embodiments. In fact, various modifications and variations can be made to the present invention without departing from the scope or spirit of the invention, which will be apparent to those skilled in the art. For example, features shown or described as part of one embodiment can be used with another embodiment to produce yet another embodiment. Therefore, it is intended that such modifications and variations be included within the scope of the appended claims and their equivalents.

[0037] like Figures 1-2As shown, the aero-engine turbine blade film cooling structure of the present invention has a hollow turbine blade, the internal cavity of which is filled with low-temperature cooling high-pressure gas introduced from the outside. Multiple pairs of double C-shaped expansion film cooling holes communicating with the internal cavity of the blade are arranged on the blade wall substrate 1. The outside of the turbine blade is a high-temperature environment formed by the high-temperature main combustion flow. The film cooling holes are used to spray the high-pressure gas in the internal cavity of the blade into the high-temperature environment in the form of a low-temperature cooling jet, and form a cooling film covering the outer wall of the turbine blade.

[0038] Each pair of C-shaped expansion film cooling holes includes an upstream C-shaped expansion film cooling hole 2 and a downstream C-shaped expansion film cooling hole 3. The upstream C-shaped expansion film cooling hole 2 and the downstream C-shaped expansion film cooling hole 3 are arranged in a front-to-back manner along the mainstream direction of the high-temperature main combustion flow. The line connecting their axes is parallel to the mainstream direction of the high-temperature main combustion flow. The inlet ends of the upstream C-shaped expansion film cooling hole 2 and the downstream C-shaped expansion film cooling hole 3 are both formed on the inner wall of the turbine blade wall base 1, and the outlet ends are both formed on the outer wall of the turbine blade wall base 1. The hole spacing e between the inlet ends of the upstream C-shaped expansion film cooling hole 2 and the downstream C-shaped expansion film cooling hole 3 is equivalent to the hole spacing f between their outlet ends.

[0039] Both the upstream C-shaped expansion film orifice 2 and the downstream C-shaped expansion film orifice 3 have a cross-section that is a semi-elliptical arc-shaped slit orifice. The major axis of the semi-elliptical arc-shaped slit orifice is basically perpendicular to the mainstream direction of the high-temperature main gas flow, and the minor axis is basically parallel to the mainstream direction of the high-temperature main gas flow. The C-shaped orifice of both the upstream C-shaped expansion film orifice 2 and the downstream C-shaped expansion film orifice 3 includes an inner ring sidewall 4 located radially inside and an outer ring sidewall 5 located radially outside. At the outlet end, the major axis radius of the outer ring sidewall 5 is a1 and the minor axis radius is b1, and the major axis radius of the inner ring sidewall 4 is a2 and the minor axis radius is b2. At the inlet end, the major axis radius of the outer ring sidewall 5 is a3 and the minor axis radius is b3, and the major axis radius of the inner ring sidewall 4 is a4 and the minor axis radius is b4. Wherein, a4≤a2, b4≤b2, a3≤a1, and b3≤b1.

[0040] The inner ring sidewall 4 and outer ring sidewall 5 of the upstream C-shaped expanding film air hole 2 and the downstream C-shaped expanding film air hole 3 are both expanding with an expansion angle α from their air inlet end to their air outlet end; the upstream C-shaped expanding film air hole 2 and the downstream C-shaped expanding film air hole 3 are both set on the blade wall substrate 1 in an inclined state, and the angle θ between their center line and the tangent direction of the blade wall substrate 1 is an acute angle.

[0041] In a preferred embodiment of the present invention, pairs of double C-shaped expanding film cooling holes are arranged in an array on the blade wall substrate 1. The hole spacing P between adjacent pairs of double C-shaped expanding film cooling holes in the direction perpendicular to the main flow of the high-temperature main combustion gas is between 3a1 and 6a1. The angle θ between the centerline of the upstream C-shaped expanding film cooling hole 2 and the downstream C-shaped expanding film cooling hole 3 and the tangent direction of the blade wall substrate 1 is between 20° and 60°. The expansion angle α between the inner ring sidewall 4 and the outer ring sidewall 5 of the upstream C-shaped expanding film cooling hole 2 and the downstream C-shaped expanding film cooling hole 3 from their inlet end to their outlet end is between 0° and 6°.

[0042] At the outlet end, the major axis radius a1 of the outer ring sidewall ranges from 4 to 10 mm, and the minor axis radius b1 ranges from 2 to 5 mm; the major axis radius a2 of the inner ring sidewall ranges from 3 to 7 mm, and the minor axis radius b2 ranges from 1 to 2 mm. At the inlet end, the major axis radius a3 of the outer ring sidewall ranges from 1.6 to 4 mm, and the minor axis radius b3 ranges from 0.4 to 2 mm; the major axis radius a4 of the inner ring sidewall ranges from 1.2 to 3 mm, and the minor axis radius b4 ranges from 0.4 to 1 mm. The hole spacing e between the inlet ends of the upstream C-shaped expanding film orifice and the downstream C-shaped expanding film orifice ranges from a1 to 2a1 mm, and the hole spacing f between their outlet ends ranges from a1 to 2a1 mm.

[0043] More specifically, Figure 1 The image shows a top view (xz plane) of the air film cooling structure for aero-engine turbine blades with a pair of double C-shaped expanding film cooling holes according to the present invention. The dashed lines represent the internal structure of the holes. It can be seen that the double C-shaped expanding film cooling holes include an upstream C-shaped expanding film cooling hole 2 and a downstream C-shaped expanding film cooling hole 3. The two holes 2 and 3 are arranged one after the other along the mainstream direction of the high-temperature main combustion gas flow. The hole spacing at the inlet end is e, and the hole spacing at the outlet end is f. Each C-shaped hole 2 and 3 includes an inner ring sidewall 4 and an outer ring sidewall 5. The major axis radius of the inner ring sidewall 4 at the inlet end is a4, and the minor axis radius is b4; the major axis radius of the outer ring sidewall 5 at the inlet end is a3, and the minor axis radius is b3; the major axis radius of the inner ring sidewall 4 at the outlet end is a2, and the minor axis radius is b2; the major axis radius of the outer ring sidewall 5 at the outlet end is a1, and the minor axis radius is b1; where a4≤a2, b4≤b2, a3≤a1, and b3≤b1; multiple rows of double C-shaped expansion film perforations can be arranged on the turbine blade wall substrate 1, and the spacing between adjacent perforation pairs in the same row is P.

[0044] Figure 2The figure shown is a cross-sectional view (yz section) of the double C-shaped expanding film air hole pair in this invention. The angle between the center line of the upper and lower C-shaped expanding film air holes 2 and 3 in the double C-shaped expanding film air hole pair and the tangent direction of the blade wall substrate 1 is θ (20~60°). The upper and lower C-shaped expanding film air holes 2 and 3 in the double C-shaped expanding film air hole pair are expanded from the air inlet end to the air outlet end, and the expansion angle is α (0~6°).

[0045] Figure 3 The existing cylindrical air film jet and the mainstream form a kidney-shaped vortex on the xy section. This vortex makes it very easy for the air film to detach from the wall, weakening the cooling effect. Figure 4 This invention utilizes double C-shaped expanding air film pores to form kidney-shaped and anti-kidney-shaped vortices. (Comparison) Figure 3 , 4 It can be seen that the improved structure of the air film pore brings several obvious technical benefits:

[0046] Existing turbine blade film cooling structures use cylindrical film cooling holes, which easily form kidney-shaped vortices near the hole outlets on the blade surface. This leads to technical problems such as cold gas lifting, reduced film coverage area, and deterioration of film cooling effect. Figure 3 As shown, the turbine blade film cooling structure proposed in this invention replaces the existing cylindrical film cooling holes with double C-shaped expansion film cooling holes. On the one hand, the cross-section of the upstream and downstream C-shaped expansion film cooling holes in the double C-shaped expansion film cooling holes is a semi-elliptical arc-shaped slit hole, and the major axis of the semi-elliptical arc-shaped slit hole is basically perpendicular to the mainstream direction of the high-temperature main gas flow, and the minor axis is basically parallel to the mainstream direction of the high-temperature main gas flow. Since the size of the double C-shaped expansion film cooling holes in the direction perpendicular to the mainstream direction is larger than the size in the direction parallel to the mainstream direction, the lateral coverage width and coverage area of ​​the film formed on the surface of the turbine blade to be cooled are increased, which improves the film cooling effect to a certain extent.

[0047] On the other hand, and more importantly, replacing the existing cylindrical film cooling orifices with double C-shaped expansion film cooling orifices results in a significant increase in pressure. Since both the upstream and downstream C-shaped expansion orifices in the double C-shaped orifices have an expansion angle α (0–6°) at their outlets, the flow velocity decreases and the pressure increases as the cryogenic cooling flow passes through the outlet with the expansion angle α. This creates a pair of vortices near the orifice outlets on the blade surface, opposite in direction to the kidney-shaped vortex rotation—the anti-kidney vortex. The anti-kidney vortex can counteract or partially counteract the effect of the kidney-shaped vortex because they have opposite directions and velocities. This reduces the longitudinal height and increases the lateral width of the kidney-shaped vortex, suppressing the adverse effects of the kidney-shaped vortex, such as gas flow instability, uneven pressure distribution in the film cooling system, and reduced film cooling efficiency. This also greatly increases the coverage area of ​​the film cooling system on the turbine blade surface to be cooled. Furthermore, the generation of the anti-kidney vortex helps reduce the variation range of the film cooling flow rate, thereby improving the uniformity and consistency of film cooling. Because the formation of the inverted kidney vortex can balance the momentum and mass transfer in the gas flow, reducing the range of gas flow variation, thereby reducing the inhomogeneity and fluctuation of the gas film. It should also be noted that, in this invention, by adjusting the hole spacing e at the inlet end and the hole spacing f at the outlet end of the upstream and downstream C-shaped expansion gas film orifices, the scale of the inverted kidney vortex can be easily adjusted, optimizing the gas flow distribution and velocity, thereby controlling the gas film coverage and cooling effect, and improving the efficiency, uniformity, and stability of gas film cooling.

[0048] Finally, it should be noted that the film cooling structure for aero-engine turbine blades with double C-shaped expanding film cooling holes provided by this invention has been verified through numerical simulations of film cooling characteristics under different parameters. Compared with simple cylindrical film cooling holes, the kidney vortex is significantly suppressed, and the film adhesion ability is greatly improved. Typical operating condition analysis shows that the cooling efficiency is improved by an average of 20%. Furthermore, this structure is simple, has wide applicability, and is a promising new cooling structure.

[0049] The objectives of this invention have been fully and effectively achieved through the above embodiments. All equivalent or simple variations made to the structures, features, and principles described in this patent concept are included within the protection scope of this patent. Those skilled in the art can make various modifications or additions to the described specific embodiments or use similar methods to replace them, as long as they do not deviate from the structure of this invention or exceed the scope defined in these claims, all of which should fall within the protection scope of this invention.

Claims

1. A film cooling structure for an aero-engine turbine blade, wherein the turbine blade is a hollow blade, its internal cavity is filled with low-temperature cooling high-pressure gas introduced from the outside, and a plurality of double-C-shaped expansion film cooling holes communicating with the internal cavity are arranged on the blade wall substrate; the outside of the turbine blade is a high-temperature environment formed by the high-temperature main combustion flow; the film cooling holes are used to inject the high-pressure gas in the internal cavity of the blade into the high-temperature environment in the form of a low-temperature cooling jet, and form a cooling film covering the outer wall of the turbine blade, characterized in that... Each of the aforementioned double C-shaped expansion film cooling hole pairs includes an upstream C-shaped expansion film cooling hole and a downstream C-shaped expansion film cooling hole. The upstream and downstream C-shaped expansion film cooling holes are arranged in a front-to-back manner along the mainstream direction of the high-temperature main combustion flow. The line connecting their axes is parallel to the mainstream direction of the high-temperature main combustion flow. The inlet ends of the upstream and downstream C-shaped expansion film cooling holes are both formed on the inner wall surface of the turbine blade wall substrate, and the outlet ends are both formed on the outer wall surface of the turbine blade wall substrate. The hole spacing e between the inlet ends of the upstream and downstream C-shaped expansion film cooling holes is equivalent to the hole spacing f between their outlet ends. The cross-sections of the upstream C-shaped expansion film orifice and the downstream C-shaped expansion film orifice are both semi-elliptical arc-shaped slits, and the major axis of the semi-elliptical arc-shaped slit is perpendicular to the mainstream direction of the high-temperature main gas flow, while the minor axis is parallel to the mainstream direction of the high-temperature main gas flow. Both the upstream and downstream C-shaped expansion film orifices include an inner ring sidewall located radially inward and an outer ring sidewall located radially outward. At the outlet end, the major axis radius of the outer ring sidewall is a1 and the minor axis radius is b1, while the major axis radius of the inner ring sidewall is a2 and the minor axis radius is b2. At the inlet end, the major axis radius of the outer ring sidewall is a3 and the minor axis radius is b3, while the major axis radius of the inner ring sidewall is a4 and the minor axis radius is b4. Furthermore, a4 ≤ a2, b4 ≤ b2, a3 ≤ a1, and b3 ≤ b1. The inner and outer ring sidewalls of the upstream and downstream C-shaped expanding air film holes both form an expansion angle from their inlet to outlet ends. α Expansion type; Both the upstream and downstream C-shaped expanding film air holes are inclinedly arranged on the blade wall substrate, with the angle between their centerlines and the tangent direction of the blade wall substrate wall. θ It is an acute angle; At the outlet end, the major axis radius a1 of the outer ring sidewall is in the range of 4 to 10 mm and the minor axis radius b1 is in the range of 2 to 5 mm, while the major axis radius a2 of the inner ring sidewall is in the range of 3 to 7 mm and the minor axis radius b2 is in the range of 1 to 2 mm. At the intake end, the major axis radius a3 of the outer ring sidewall ranges from 1.6 to 4 mm, and the minor axis radius b3 ranges from 0.4 to 2 mm; the major axis radius a4 of the inner ring sidewall ranges from 1.2 to 3 mm, and the minor axis radius b4 ranges from 0.4 to 1 mm. The numerical range of the hole spacing e between the inlet ends of the upstream C-shaped expansion air film hole and the downstream C-shaped expansion air film hole is between a1 and 2a1, and the numerical range of the hole spacing f between their outlet ends is between a1 and 2a1.

2. The air film cooling structure for aero-engine turbine blades according to claim 1, characterized in that, The pairs of double C-shaped expansion film cooling holes are arranged in an array on the blade wall substrate, and the hole spacing P between two adjacent pairs of double C-shaped expansion film cooling holes in the direction perpendicular to the main flow of the high-temperature main gas flow is between 3a1 and 6a1.

3. The air film cooling structure for aero-engine turbine blades according to claim 1, characterized in that, The angle between the centerlines of the upstream and downstream C-shaped expanding film vents and the tangent direction of the blade wall substrate. θ The numerical range is between 20 and 60°.

4. The air film cooling structure for aero-engine turbine blades according to claim 1, characterized in that, The expansion angle between the inner and outer ring sidewalls of the upstream and downstream C-shaped expanding film orifices from their inlet to outlet ends. α The numerical range is between 0 and 6°.

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