A turbine blade film cooling structure of a three-hole combination of middle expansion and both sides contraction

Abstract

The present application relates to a kind of intermediate expansion two sides contract three-hole combination turbine blade film cooling structure, to improve the cooling efficiency of turbine blade and reduce aerodynamic loss.The structure includes an intermediate film jet channel and the first side film jet channel and the second side film jet channel in the spanwise of blade on its two sides, the expansion design of intermediate film jet channel helps to reduce jet velocity, increase outlet pressure, so as to effectively inhibit the penetration ability of high temperature main stream, enhance the stability of film.Two sides film jet channel is contracted design, and the formation of reverse kidney-shaped vortex is promoted by lifting jet velocity, reducing pressure, effectively inhibiting the negative effect of traditional kidney-shaped vortex, enhancing the mixing effect of film and main stream, so as to improve the film cooling efficiency.The structural characteristics and principle of the present application are strong, with good adjustability, can be widely applied to turbine guide vane, moving blade, combustion chamber and other high-temperature components cooling, with wide application prospect.

Description

Technical Field

[0001] This invention belongs to the field of aero-engine turbine blade cooling technology, and relates to the design and optimization of turbine blade film cooling holes. Specifically, it relates to a turbine blade film cooling structure with a three-hole combination of central expansion and two-sided contraction, which can effectively improve the film cooling effect of turbine blades and reduce aerodynamic losses. Background Technology

[0002] Aero engines are the power source of aircraft, and their performance improvements directly affect flight performance and safety. To improve the combustion efficiency and thrust of aero engines, it is necessary to increase the turbine inlet temperature, thereby improving cycle efficiency. However, the increased turbine inlet temperature also presents a significant challenge to turbine blade cooling. Turbine blades are one of the core components of aero engines, operating in extremely harsh environments, enduring multiple loads including high temperature, high pressure, and high-speed rotation. The turbine inlet temperature of advanced civil high-bypass aero engines has exceeded 2000K, a temperature far exceeding the maximum temperature resistance limit that turbine blade materials can withstand. Under such extreme operating conditions, without effective cooling measures, turbine blades are prone to creep or damage due to high temperatures, leading to decreased engine performance or failure.

[0003] To protect the service life of turbine blades at high temperatures, film cooling (FSR) technology is widely used. FSR utilizes low-temperature cooling gas introduced from inside the blade to form a film on the blade surface, blocking the heat radiation and convection of the high-temperature main combustion flow, thereby reducing the blade surface temperature. The advantages of FSR are high cooling efficiency, low cooling gas consumption, and minimal impact on engine performance. The efficiency of FSR directly affects the cooling effect on turbine blades and the overall performance of the engine. FSR efficiency is influenced not only by flow parameters such as air-to-air ratio, density ratio, and mainstream Reynolds number, but also by the geometric parameters of the film cooling orifice.

[0004] Currently, researchers have conducted extensive research and achieved certain results on the shape of film cooling (WCC) orifices, including composite-angle WCCs (WCCs with certain inclination angles in both the axial and circumferential directions) and irregularly shaped orifices (WCCs where the outlet area differs from the inlet area, such as expansion orifices, contraction orifices, and fan-shaped orifices). The results show that improving the shape of WCCs can effectively enhance the film cooling effect compared to traditional cylindrical WCCs. This is mainly because during the mixing process of the jet from a cylindrical WCC with the mainstream, the temperature and velocity difference between the jet and the mainstream creates a kidney-shaped vortex. This kidney-shaped vortex forces the low-temperature cooling jet to detach from the wall, while the high-temperature mainstream gas flow re-attaches to the wall, resulting in a reduction in the film coverage area and a decrease in the film cooling effect. Improving the shape of the WCC, to a certain extent, increases the outlet area, reduces the jet outlet velocity, and weakens the influence of the kidney-shaped vortex, thereby achieving the goal of improving the film cooling effect. For example, composite-angle WCCs can bring the jet closer to the wall, increasing the film coverage area; irregularly shaped orifices can change the jet diffusion angle, increasing the film thickness and uniformity.

[0005] However, improving the shape of film cooling holes also presents some problems and challenges. First, the improved film cooling hole structure, such as irregularly shaped holes, is relatively complex, difficult to manufacture, and costly. This is especially true for areas with high curvature, such as the leading and trailing edges of the blades, where machining the film cooling holes is even more challenging. Simultaneously, it has a certain impact on the mainstream aerodynamic performance of the blade surface and may increase aerodynamic losses, thus affecting the overall engine efficiency. For example, composite-angle film cooling holes can generate certain lateral forces on the blade surface, affecting blade stability. Furthermore, irregularly shaped holes can alter the surface roughness of the blade, affecting drag and lift.

[0006] In conclusion, although current film cooling technology has made some progress, how to simplify the structure of the film cooling holes, reduce manufacturing difficulty and cost, and minimize the impact on blade aerodynamic performance while ensuring film cooling efficiency remains a pressing technical challenge. Solving these problems will have a significant impact on improving the performance and reliability of aero-engines. Summary of the Invention

[0007] (I) Purpose of the Invention

[0008] To address the limitations of existing film cooling technologies, particularly the reniform vortices induced by traditional cylindrical film cooling orifices and the resulting reduction in cooling efficiency and increased aerodynamic losses, this invention proposes a turbine blade film cooling structure with a centrally expanding and laterally contracting three-orifice design. By utilizing fluid dynamics theory, the vortex structure after mixing of the film cooling outflow and the mainstream is reconstructed based on velocity and pressure distribution, thereby improving film cooling efficiency and reducing aerodynamic losses. The core of this invention lies in the three-orifice design with a centrally expanding and laterally contracting orifice. The expansion of the central orifice helps reduce jet velocity and increase outlet pressure, effectively suppressing the penetration ability of the high-temperature mainstream and enhancing film cooling stability. The contracting design of the two side orifices, by increasing jet velocity and reducing pressure, promotes the formation of anti-reniform vortices, effectively suppressing the negative effects of traditional reniform vortices and enhancing the mixing effect between the film cooling and the mainstream, thus improving film cooling efficiency. The turbine blade film cooling structure of the present invention, which combines a central expansion and two side contractions with three holes, has the advantages of unique structure and strong functionality. It significantly improves the film cooling effect, effectively enhances the cooling performance of turbine blades, extends their service life, and has broad application prospects, especially in the field of cooling technology for high-performance aero-engine turbine blades.

[0009] (II) Technical Solution

[0010] To achieve the objective of this invention and solve its technical problems, the present invention adopts the following technical solution:

[0011] A turbine blade film cooling structure with a central expansion and two side contractions, designed to improve the cooling performance of turbine blades and reduce aerodynamic losses, comprises multiple turbine blades evenly distributed circumferentially, with a main gas flow channel formed between adjacent turbine blades, and each turbine blade having a hollow cavity communicating with an external cooling gas source. Its distinguishing feature is that...

[0012] Each turbine blade has a plurality of film-forming hole assemblies distributed along the blade span from the blade root to the blade tip on its base surface of the suction and / or pressure surfaces. Each film-forming hole assembly consists of three film-forming jet channels arranged chordally aligned and adjacent to each other along the blade span: a central film-forming jet channel, and a first side film-forming jet channel and a second side film-forming jet channel symmetrically distributed on both sides of the central film-forming jet channel along the blade span.

[0013] The inlet ports of the intermediate film jet channel, the first side film jet channel, and the second side film jet channel are all formed on the inner side of the turbine blade substrate surface and communicate with the hollow cavity of the turbine blade filled with cooling gas. The outlet ports are all formed on the outer side of the turbine blade substrate surface and communicate with the main combustion gas flow channel. Each inlet port is formed as a circular hole, and each outlet port is formed as an irregularly shaped hole. Furthermore, the three irregularly shaped outlet ports of each film jet hole assembly are arranged in a crescent shape along the blade spanwise on the turbine blade substrate surface.

[0014] The intermediate film jet channel is formed in a gradually expanding shape from its inlet to its outlet. The projection of the channel's centerline onto the turbine blade substrate surface extends in the same direction as the blade's chord. Furthermore, the area of ​​its irregularly shaped outlet port is larger than the area of ​​its circular inlet port. This reduces the film jet velocity at the outlet and increases the film jet outlet pressure, thereby facilitating the formation of a stable low-speed, high-pressure cooling film region and suppressing the penetration capability of the high-temperature main combustion gas flow.

[0015] The first and second side air film jet channels are formed in a gradually narrowing shape from their inlet to their outlet. The area of ​​the outlet port, which has an irregular hole, is smaller than the area of ​​the inlet port, which has a circular hole. This increases the air film jet velocity at the outlet and reduces the air film jet outlet pressure. This facilitates the formation of a high-speed, low-pressure region on both sides of the low-speed, high-pressure region of the intermediate air film jet channel, promotes the formation of an anti-kidney vortex, and weakens the negative effects of the kidney vortex. This enhances the air film adhesion effect and improves the air film cooling capacity.

[0016] Preferably, in each of the air film jet assemblies, the irregular holes at the outlet ports of the intermediate air film jet channel, the first side air film jet channel, and the second side air film jet channel each include an upstream elliptical arc located upstream in the blade chord direction, a downstream elliptical arc located downstream in the blade chord direction, and two transition arcs for connecting the upstream and downstream elliptical arcs and located at the two ends of the elliptical arcs respectively in the blade span direction. The major axis directions of the upstream and downstream elliptical arcs are consistent with the blade chord direction, and the minor axis directions of the upstream and downstream elliptical arcs are consistent with the blade span direction.

[0017] Furthermore, in each of the aforementioned air film jet assemblies, the circular holes at the air inlet ports of the intermediate air film jet channel, the first side air film jet channel, and the second side air film jet channel have the same diameter D, with the diameter D ranging from 0.6 mm to 4 mm; the major axis radius of the upstream elliptical arc is a1, and the minor axis radius is b1, with a1 ranging from 4D to 6D and b1 ranging from 2D to 4D; the major axis radius of the downstream elliptical arc is a2, and the minor axis radius is b2, with a2 ranging from 4D to 6D and b2 ranging from 2D to 4D; the radius of the transition arc is r, with r ranging from 0.1D to 0.2D.

[0018] Preferably, in each of the air film jet assembly, the circular holes at the air inlet ports of the intermediate air film jet channel, the first side air film jet channel, and the second side air film jet channel have the same chordal position, and the hole spacing L between the centers of two adjacent circular holes ranges from 1.8D to 3D. In the irregular holes at the air outlet ports of the intermediate air film jet channel, the first side air film jet channel, and the second side air film jet channel, the distance t between the adjacent edges of two adjacent irregular holes ranges from 0.16D to 0.25D, where D is the diameter of the circular hole.

[0019] Preferably, in each of the film jet assemblies, the angle between the centerline of each film jet channel and the surface of the turbine blade substrate in the chord direction of the blade is θ. The angle θ of the jet channel is in the range of 20° to 60° and is optimized based on the aerodynamic characteristics and cooling requirements of the turbine blade substrate surface to ensure that the film jet can effectively cover the blade surface and form a stable cooling film.

[0020] Preferably, in each of the gas film pore assemblies, the expansion shape design of the intermediate gas film jet channel adopts a gradually increasing diameter ratio, so that the diameter of the channel gradually increases from the inlet port to the outlet port and the rate of increase gradually accelerates, thereby forming a more stable flow velocity distribution in the intermediate gas film jet channel, so as to optimize the dynamic characteristics of the gas film jet, ensure that the formation of the low-speed high-pressure region is more stable, and suppress the penetration of high-temperature mainstream gas.

[0021] Preferably, in each of the air film jet assemblies, the contraction shape design of the first side air film jet channel and the second side air film jet channel adopts a linear contraction method, so that the channel diameter gradually decreases from the air inlet port to the air outlet port at a constant ratio, thereby ensuring that the airflow gradually increases in speed when passing through the side air film jet channel and forms a stable high-speed low-pressure jet at the air outlet port. Meanwhile, strong shearing effects are generated on both sides of the low-speed high-pressure region formed by the middle air film jet channel to promote the formation of anti-kidney vortex and inhibit the formation of kidney vortex, thereby enhancing the air film adhesion effect and improving the air film cooling capacity.

[0022] Furthermore, in the assembly of several exhaust film holes distributed along the blade span, the contraction ratio of the first side air film jet channel and the second side air film jet channel is adjusted according to the thermal load and aerodynamic characteristics of different spanwise parts of the turbine blade, so as to adapt to the working environment of the turbine blade, improve the air film cooling effect and reduce aerodynamic losses.

[0023] Furthermore, at the blade root, the contraction ratio of the first-side film jet channel and the second-side film jet channel is set to be relatively small to maintain a relatively high outlet pressure of the film jet to suppress the penetration ability of the high-temperature mainstream, while avoiding an excessively large contraction ratio that would lead to excessively high film jet velocity and increase aerodynamic losses. At the blade tip, a larger contraction ratio is set to increase the outlet velocity of the film jet and reduce the outlet pressure of the film jet, thereby facilitating the formation of a high-speed, low-pressure cooling film region and promoting the mixing of the film jet with the mainstream, improving the cooling effect of the film jet, and optimizing the dynamic characteristics of the film jet to reduce aerodynamic losses.

[0024] Preferably, each turbine blade has a plurality of exhaust film hole assemblies distributed along the blade chord direction on the base surface of the suction surface and / or pressure surface. Each exhaust film hole assembly includes multiple air film hole assemblies distributed along the blade spanwise from the blade root to the blade tip. The hole spacing P between two adjacent air film hole assemblies in the same row in the blade spanwise direction is defined as the distance between the center lines of the intermediate air film jet channels of two adjacent air film hole assemblies in the same row in the blade spanwise direction. The range of P is between 7D and 12D, where D is the diameter of the circular hole.

[0025] (III) Technical Effects

[0026] Compared with the prior art, the turbine blade film cooling structure of the present invention, which combines a central expansion and two side contractions with three holes, has the following beneficial and significant technical effects:

[0027] (1) The turbine blade film cooling structure of the present invention, which is a combination of three holes with a middle expansion and two sides contraction, is composed of three film jet channels, namely a middle film jet channel and a first side film jet channel and a second side film jet channel located on both sides. The inlet end of the three film jet channels is set as a conventional cylindrical hole, while the outlet end is set as an irregular hole arranged in a crescent shape along the blade span. The structure features are prominent. The middle film jet channel is expansion-shaped, with a reduced outlet velocity and increased pressure. The two side film jet channels are contraction-shaped, with a higher outlet velocity and reduced pressure. This is conducive to inducing the generation of anti-kidney vortex and inhibiting the formation of kidney vortex, thereby enhancing the adhesion of the film to the wall.

[0028] (2) The turbine blade air film cooling structure of the present invention, which combines a three-hole combination of central expansion and two-side contraction, forms a low-speed, high-pressure zone at the outlet of the central hole and a high-speed, low-pressure zone at the outlets of the two side holes. It makes full use of the principles of fluid mechanics and realizes the reconstruction of the vortex structure after the air film outflow and the mainstream are mixed. The expansion design of the central air film jet channel helps to reduce the jet velocity and increase the pressure at the outlet, thereby effectively suppressing the penetration ability of the high-temperature mainstream and enhancing the stability of the air film. The contraction design of the two side air film jet channels promotes the formation of the anti-kidney vortex by increasing the jet velocity and reducing the pressure, thereby effectively suppressing the negative effects of the traditional kidney vortex. In the end, the overall air film cooling effect is improved and the aerodynamic loss is reduced.

[0029] (3) The turbine blade film cooling structure of the present invention, which combines a central expansion and two side contractions with three holes, can flexibly adjust the film outflow structure and control the development of the kidney vortex by adjusting the radius of different arc segments according to different application requirements and working conditions, thus facilitating the achievement of good cooling effects for different applications. In addition, the present invention can optimize the design of parameters such as the shape, size, angle, and spacing of the film cooling holes according to the thermal load and aerodynamic characteristics of different parts of the turbine blade, thereby achieving customized film cooling and improving the adaptability and sensitivity of cooling.

[0030] (4) The present invention has a wide range of applications. This structure 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. This structure can not only improve the cooling effect of turbine blades and extend their service life, but also reduce aerodynamic losses and improve turbine performance. It has broad application prospects, especially in the field of cooling technology for high-performance aero-engine turbine blades, where it has significant meaning and value.

[0031] (5) The present invention significantly improves cooling efficiency. This structure fundamentally suppresses the negative effects of the kidney-shaped vortex, resulting in a significant improvement in cooling performance. The kidney-shaped vortex is an unfavorable factor in film cooling, as it disrupts the continuity and stability of the film, leading to increased heat load on the blade surface and reduced cooling efficiency. The present invention, through a three-hole combination design with expansion in the middle and contraction on both sides, effectively suppresses the development of the kidney-shaped vortex, promotes the formation of the anti-kidney-shaped vortex, and enhances the adhesion effect of the film, thereby improving the film cooling efficiency and enhancing the cooling performance of the blade. Attached Figure Description

[0032] Figure 1 The image shown is a top view (xz plane) of the turbine blade film cooling structure with a combination of three holes for expansion in the middle and contraction on both sides according to the present invention, where x is the blade spanwise and z is the blade chordwise.

[0033] Figure 2The figure shown is a cross-sectional view (yz section) of the turbine blade film cooling structure with a combination of three holes for expansion in the middle and contraction on both sides according to the present invention, where y is the blade thickness direction and z is the blade chord direction.

[0034] Figure 3 The diagram shows the flow structure of the turbine blade film cooling structure of the present invention, which is a combination of three holes with a central expansion and two sides with contraction, after the jet is ejected.

[0035] Explanation of reference numerals in the attached figures:

[0036] 1- Turbine blade substrate surface, 2- Film gas flow port assembly, 3- First side film gas flow channel, 4- Middle film gas flow channel, 5- Second side film gas flow channel, 6- Upstream elliptical arc, 7- Downstream elliptical arc, 8- Transition arc, 9- Inlet port, 10- Outlet port, 11- Main combustion gas flow, 12- Film gas flow, 13- Main combustion gas flow direction, 14- Blade spanwise, 15- High-speed low-pressure zone at both side hole outlets, 16- Low-speed high-pressure zone at the middle hole outlet, 17- Kidney-shaped vortex, 18- Anti-kidney-shaped vortex. Detailed Implementation

[0037] 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.

[0038] Figure 1 The image shown is a top view (xz plane) of the turbine blade film cooling structure of the present invention, which features a combination of three holes: one for central expansion and two for side contraction. Specifically, as... Figures 1-3As shown, the turbine blade film cooling structure of the present invention, which is a combination of three holes with a central expansion and two side contractions, includes multiple turbine blades evenly distributed in the circumferential direction. A main gas flow channel is formed between adjacent turbine blades. Each turbine blade is provided with a hollow cavity that communicates with an external cooling gas source. Multiple film hole assemblies 2 are provided on the base surface 1 of the suction surface and / or pressure surface of each turbine blade, distributed from the blade root to the blade tip along the blade span 14. Each film hole assembly 2 is composed of three film jet channels that are chordally aligned and arranged adjacent to each other in the blade span 14. These are a central film jet channel 4 and a first side film jet channel 3 and a second side film jet channel 5 that are symmetrically distributed on both sides of the central film jet channel in the blade span direction. The air inlet ports 9 of the intermediate film jet channel 4, the first side film jet channel 3 and the second side film jet channel 5 are all formed on the inner side of the turbine blade substrate surface 1 and are connected to the hollow cavity of the turbine blade filled with cooling gas. The air outlet ports 10 are all formed on the outer side of the turbine blade substrate surface 1 and are connected to the main combustion gas flow channel. Each air inlet port 9 is formed as a circular hole and each air outlet port 10 is formed as an irregular hole. The three air outlet ports 9 with irregular holes of each film jet assembly 2 are arranged in a crescent shape along the blade span on the turbine blade substrate surface 1.

[0039] Furthermore, the intermediate film jet channel 4 is formed in a gradually expanding shape from its inlet port 9 to its outlet port 10. The projection of the channel centerline onto the turbine blade substrate surface 1 extends in the same direction as the blade chord. Moreover, the area of ​​the outlet port 10, which has an irregular hole, is larger than the area of ​​the inlet port 9, which has a circular hole. This reduces the film jet velocity at the outlet and increases the film jet outlet pressure, thereby facilitating the formation of a stable low-speed, high-pressure cooling film region and suppressing the penetration ability of the high-temperature main combustion flow. The first side air film jet channel 3 and the second side air film jet channel 5 are formed in a gradually narrowing shape from their inlet port 9 to their outlet port 10. The area of ​​the outlet port 10, which has an irregular hole, is smaller than the area of ​​the inlet port 9, which has a circular hole. This increases the air film jet velocity at the outlet and reduces the air film jet outlet pressure. This facilitates the formation of a high-speed, low-pressure region 16 on both sides of the low-speed, high-pressure region 10 in the middle air film jet channel 4. This promotes the formation of an anti-kidney vortex 18 and weakens the negative effect of the kidney vortex 17, thereby enhancing the air film adhesion effect and improving the air film cooling capacity.

[0040] from Figure 1As can be seen, the structure includes a central film jet channel 4, a first-side film jet channel 3, and a second-side film jet channel 5. The central film jet channel 4 is expansion-shaped, while the first-side film jet channel 3 and the second-side film jet channel 5 are contraction-shaped and symmetrical about the central film jet channel 4. The inlet ports of the three film jet channels are three circular holes, which are systematically positioned along the z-axis of the blade. The diameter of the circular holes is D, which ranges from 0.6 to 4 mm. The outlet ports of the three film jet channels are three irregularly shaped holes, which are arranged in a crescent shape along the x-axis of the blade on the surface 1 of the turbine blade substrate.

[0041] from Figure 1 It can also be seen that the irregularly shaped holes at the outlet ports of the three film jet channels are composed of elliptical arcs and circular arcs. These include an upstream elliptical arc located upstream in the blade chord z direction, a downstream elliptical arc located downstream in the blade chord z direction, and two transition circular arcs connecting the upstream and downstream elliptical arcs and located at the ends of the elliptical arcs in the blade spanwise x direction. The major axes of both the upstream and downstream elliptical arcs are aligned with the blade chord z direction, and their minor axes are aligned with the blade spanwise x direction. The major axis radius of the upstream elliptical arc is a1, and the minor axis radius is b1, with a1 ranging from 4D to 6D and b1 ranging from 2D to 4D. The major axis radius of the downstream elliptical arc is a2, and the minor axis radius is b2, with a2 ranging from 4D to 6D and b2 ranging from 2D to 4D. A transition circular arc with radius r connects the upstream and downstream elliptical arcs, with r ranging from 0.1D to 0.2D.

[0042] In addition, in each air film orifice assembly, the circular holes at the air inlet ports of the three air film jet channels have the same spanwise position, the hole spacing between the three circular holes is L, and L ranges from 1.8D to 3D; the distance between the edges of the outlet orifices of the three air film jet channels is t, and t ranges from 0.16D to 0.25D.

[0043] In a preferred embodiment of the present invention, multiple air film perforation assemblies can be arranged along the blade chord on the substrate of the suction surface and / or pressure surface of each turbine blade. Each air film perforation assembly includes multiple air film perforation assemblies distributed along the blade spanwise from the blade root to the blade tip. The hole spacing P between two adjacent air film perforation assemblies in the same row in the blade spanwise is defined as the distance between the center lines of the intermediate air film jet channels of two adjacent air film perforation assemblies in the same row in the blade spanwise. The range of P is between 7D and 12D, where D is the diameter of the circular hole.

[0044] Figure 2The diagram shows a cross-sectional view (yz section) of a three-hole combined air film cooling structure with central expansion and two-sided contraction. The angle between the center line of the holes and the horizontal direction is θ, which is between 20° and 60°.

[0045] Figure 3 This diagram illustrates the flow structure of a three-hole combined air-film cooling structure with a central expansion and two-sided contraction. Traditional cylindrical air-film jets and the mainstream form a kidney-shaped vortex in the xy-section. This vortex easily causes the air film to detach from the wall, weakening the cooling effect. With the three-hole combined air-film cooling structure, a low-speed, high-pressure zone forms at the outlet of the central air-film jet channel 4, while high-speed, low-pressure zones form at the outlets of the left and right holes. This flow distribution induces an anti-kidney-shaped vortex, whose rotation direction is opposite to that of the kidney-shaped vortex, thus mitigating the negative effects of the kidney-shaped vortex. In this structure, both the upstream and downstream elliptical arcs expand to the sides, reducing the outflow from the central region and increasing the spanwise outflow, resulting in more uniform air-film cooling. The upstream and downstream elliptical arcs can be elliptical or other shapes. By adjusting the radii of the upstream and downstream elliptical arcs and the transition arc, the crescent-shaped outlet structure can be flexibly adjusted to control the distribution of the air film along the substrate surface, thereby controlling the air-film cooling capacity.

[0046] 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 turbine blade film cooling structure with a central expansion and two side contractions, used to improve the cooling performance of turbine blades and reduce aerodynamic losses, comprising multiple turbine blades evenly distributed in the circumferential direction, with a main gas flow channel formed between adjacent turbine blades, and each turbine blade having a hollow cavity communicating with an external cooling gas source, characterized in that, Each turbine blade has a plurality of film-forming hole assemblies distributed along the blade span from the blade root to the blade tip on its base surface of the suction and / or pressure surfaces. Each film-forming hole assembly consists of three film-forming jet channels arranged chordally aligned and adjacent to each other along the blade span: a central film-forming jet channel, and a first side film-forming jet channel and a second side film-forming jet channel symmetrically distributed on both sides of the central film-forming jet channel along the blade span. The inlet ports of the intermediate film jet channel, the first side film jet channel, and the second side film jet channel are all formed on the inner side of the turbine blade substrate surface and communicate with the hollow cavity of the turbine blade filled with cooling gas. The outlet ports are all formed on the outer side of the turbine blade substrate surface and communicate with the main combustion gas flow channel. Each inlet port is formed as a circular hole, and each outlet port is formed as an irregularly shaped hole. Furthermore, the three irregularly shaped outlet ports of each film jet hole assembly are arranged in a crescent shape along the blade spanwise on the turbine blade substrate surface. The intermediate film jet channel is formed in a gradually expanding shape from its inlet to its outlet. The projection of the channel's centerline onto the turbine blade substrate surface extends in the same direction as the blade's chord. Furthermore, the area of ​​its irregularly shaped outlet port is larger than the area of ​​its circular inlet port. This reduces the film jet velocity at the outlet and increases the film jet outlet pressure, thereby facilitating the formation of a stable low-speed, high-pressure cooling film region and suppressing the penetration capability of the high-temperature main combustion gas flow. The first and second side air film jet channels are formed in a gradually narrowing shape from their inlet to their outlet. The area of ​​the outlet port, which has an irregular hole, is smaller than the area of ​​the inlet port, which has a circular hole. This increases the air film jet velocity at the outlet and reduces the air film jet outlet pressure. This facilitates the formation of a high-speed, low-pressure region on both sides of the low-speed, high-pressure region of the intermediate air film jet channel, promotes the formation of an anti-kidney vortex, and weakens the negative effects of the kidney vortex. This enhances the air film adhesion effect and improves the air film cooling capacity.

2. The turbine blade film cooling structure with a central expansion and two side contractions combined according to claim 1, characterized in that, In each of the aforementioned film jet duct assemblies, the irregularly shaped holes at the outlet ports of the intermediate film jet channel, the first side film jet channel, and the second side film jet channel each include an upstream elliptical arc located upstream in the blade chord direction, a downstream elliptical arc located downstream in the blade chord direction, and two transition arcs connecting the upstream and downstream elliptical arcs and located at the two ends of the elliptical arcs respectively in the blade span direction. The major axis directions of the upstream and downstream elliptical arcs are consistent with the blade chord direction, and the minor axis directions of the upstream and downstream elliptical arcs are consistent with the blade span direction.

3. The turbine blade film cooling structure with a central expansion and two side contractions combined according to claim 2, characterized in that, In each of the aforementioned air film jet assemblies, the circular holes at the air inlet ports of the intermediate air film jet channel, the first side air film jet channel, and the second side air film jet channel have the same diameter D, which ranges from 0.6 mm to 4 mm; the major axis radius of the upstream elliptical arc is a1, and the minor axis radius is b1, where a1 ranges from 4D to 6D, and b1 ranges from 2D to 4D; the major axis radius of the downstream elliptical arc is a2, and the minor axis radius is b2, where a2 ranges from 4D to 6D, and b2 ranges from 2D to 4D; the radius of the transition arc is r, which ranges from 0.1D to 0.2D.

4. The turbine blade film cooling structure with a central expansion and two side contractions combined according to claim 1, characterized in that, In each of the aforementioned air film jet assemblies, the circular holes at the air inlet ports of the intermediate air film jet channel, the first side air film jet channel, and the second side air film jet channel have the same chordal position, and the hole spacing L between the centers of two adjacent circular holes ranges from 1.8D to 3D. In the irregular holes at the air outlet ports of the intermediate air film jet channel, the first side air film jet channel, and the second side air film jet channel, the distance t between the adjacent edges of two adjacent irregular holes ranges from 0.16D to 0.25D, where D is the diameter of the circular hole.

5. The turbine blade film cooling structure with a central expansion and two side contractions combined according to claim 1, characterized in that, In each of the aforementioned film jet assemblies, the angle between the centerline of each film jet channel and the surface of the turbine blade substrate in the chord direction of the blade is θ. The angle θ of the jet channel is in the range of 20° to 60° and is optimized based on the aerodynamic characteristics and cooling requirements of the turbine blade substrate surface to ensure that the film jet can effectively cover the blade surface and form a stable cooling film.

6. The turbine blade film cooling structure with a central expansion and two side contractions combined according to claim 1, characterized in that, In each of the aforementioned air film orifice assemblies, the expansion shape design of the intermediate air film jet channel adopts a gradually increasing diameter ratio, so that the diameter of the channel gradually increases from the inlet port to the outlet port and the rate of increase gradually accelerates, thereby forming a more stable flow velocity distribution in the intermediate air film jet channel, so as to optimize the dynamic characteristics of the air film jet, ensure that the formation of the low-speed high-pressure region is more stable, and suppress the penetration of high-temperature mainstream gas.

7. The turbine blade film cooling structure with a central expansion and two side contractions combined according to claim 1, characterized in that, In each of the aforementioned air film jet assemblies, the first and second side air film jet channels are designed with a linear contraction shape, so that the channel diameter gradually decreases at a constant ratio from the inlet port to the outlet port. This ensures that the airflow gradually increases in speed as it passes through the side air film jet channels and forms a stable high-speed, low-pressure jet at the outlet port. Meanwhile, strong shearing effects are generated on both sides of the low-speed, high-pressure region formed by the middle air film jet channel to promote the formation of anti-kidney vortices and inhibit the formation of kidney vortices, thereby enhancing the air film adhesion effect.

8. The turbine blade film cooling structure with a central expansion and two side contractions combined according to claim 7, characterized in that, In the assembly of several exhaust film holes distributed along the blade span, the contraction ratio of the first side air film jet channel and the second side air film jet channel is adjusted according to the thermal load and aerodynamic characteristics of different spanwise parts of the turbine blade, so as to adapt to the working environment of the turbine blade, improve the air film cooling effect and reduce aerodynamic losses.

9. The turbine blade film cooling structure with a central expansion and two side contractions combined according to claim 8, characterized in that, At the blade root, the contraction ratio of the first and second side film air jet channels is set relatively small to maintain a relatively high outlet pressure of the film air jet to suppress the penetration ability of the high-temperature mainstream, while avoiding excessive contraction ratio that would lead to excessively high film air jet velocity and increase aerodynamic losses. At the blade tip, a larger contraction ratio is set to increase the outlet velocity of the film air jet and reduce the outlet pressure of the film air jet, thereby facilitating the formation of a high-speed, low-pressure cooling film region and promoting the mixing of the film air jet with the mainstream, improving the cooling effect of the film air, while optimizing the dynamic characteristics of the film air jet and reducing aerodynamic losses.

10. The turbine blade film cooling structure with a central expansion and two side contractions combined according to claim 1, characterized in that, Each turbine blade has a plurality of exhaust film hole assemblies distributed along the blade chord direction on the base surface of the suction surface and / or pressure surface. Each exhaust film hole assembly includes multiple air film hole assemblies distributed along the blade spanwise from the blade root to the blade tip. The hole spacing P between two adjacent air film hole assemblies in the same row in the blade spanwise direction is defined as the distance between the center lines of the intermediate air film jet channels of two adjacent air film hole assemblies in the same row in the blade spanwise direction. The range of P is between 7D and 12D, where D is the diameter of the circular hole.

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