Aero-engine turbine blade gill area gradient aperture air film cooling layout structure

Abstract

The present application relates to a kind of turbine blade gill area of aero-engine turbine blade gradient aperture air film cooling layout structure, aims at the cooling demand of turbine blade under high temperature, high speed and high pressure working condition, especially in the key area of blade gill area, the curvature changes greatly, flow changes dramatically, provide improved cooling effect.The structure is by setting the gradient change of air film aperture in blade gill area along the blade span, realize gradually adjusting the size of air film hole according to the change of flow velocity and heat load distribution, to optimize the distribution and coverage effect of cooling airflow.The structure has the advantages of simple structure, strong pertinence, flexible arrangement, obvious cold effect improvement, etc., improves the cooling efficiency and cooling uniformity of blade gill area, effectively prolongs the service life of turbine blade, and improves the overall performance and reliability of engine.The present application also provides several optimized air film hole shapes, arrangement mode and inclination, to further improve the coverage, stability, adhesion and cooling effect of air film.

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 the air film cooling layout structure in the gill region of turbine blades. Specifically, it is a gradient aperture air film cooling layout structure in the gill region of aero-engine turbine blades. By gradually increasing or decreasing the air film aperture from the blade root to the blade tip according to the distribution of flow velocity along the spanwise direction in the blade gill region, a complete and good air film coverage of the gill region is achieved, thereby enhancing the cooling effect. Background Technology

[0002] Turbine blades are one of the core components of an aero-engine, operating in extremely harsh environments, subjected to high temperatures, high pressures, and high-speed airflow. Therefore, effective cooling technologies are essential to protect the structural integrity and material properties of the blades. Film cooling (FSM) is a commonly used turbine blade cooling technique. Its principle involves arranging a series of film cooling holes on the blade surface. Cool air ejected from these holes forms a film covering the blade surface, thus isolating it from the direct impact of hot airflow and reducing the blade's surface temperature and thermal stress.

[0003] In existing film cooling schemes for aero-engine turbine blades, the layout typically involves multiple rows of densely packed film cooling holes at the front edges of both the pressure and suction surfaces of the turbine blades to achieve spray cooling. The pressure surface has 3-6 rows of evenly distributed film cooling holes, while the suction surface has 2-4 rows, providing comprehensive thermal protection for the blades. The advantages of this layout are its simplicity and ease of implementation, and it can, to a certain extent, meet the cooling requirements of turbine blades, protecting them from direct attack by high-temperature exhaust gases.

[0004] However, with the pursuit of high efficiency and high performance in aero-engines, turbine inlet temperatures are constantly increasing, placing higher demands on turbine blade cooling technology. Conventional uniform film cooling technology is no longer sufficient to meet design requirements. In particular, the complex shape and geometry of turbine blades result in extremely complex flow structures. Uniformly arranged film cooling holes can easily lead to areas on the turbine blade surface not being covered by the film, causing blade ablation and affecting reliability and service life. The gill region (the transition area between the leading edge and the blade body) requires special attention. The gill region exhibits significant curvature changes, large pressure gradients, and dramatic flow variations along the flow direction. The flow velocity also changes considerably along the spanwise direction from the blade root to the tip, posing a significant challenge to the cooling layout design in this region. Therefore, it is necessary to implement targeted and refined film cooling layouts.

[0005] At present, research on the air film cooling layout of the turbine blade gill area mainly focuses on the following aspects: (1) changing the shape of the air film holes, such as using non-circular air film holes such as elliptical, fan-shaped, and rhomboid shapes, to increase the outlet area of ​​the air film holes and improve the coverage and stability of the air film; (2) changing the arrangement of the air film holes, such as using non-uniformly arranged air film holes such as staggered, interlaced, and spiral arrangements, to increase the interference effect of the air film and improve the air film's resistance to blow-off; (3) changing the inclination angle of the air film holes, such as using different inclination angles of the air film holes such as forward tilt, reverse tilt, and compound tilt, to adjust the flow direction and velocity of the air film and improve the adhesion and cooling effect of the air film.

[0006] Although the above studies have improved the film cooling performance of the turbine blade gill region to some extent, the following technical problems or challenges still exist: (1) Changing the shape of the film cooling holes will increase the manufacturing difficulty and cost of the film cooling holes, and the influence of different shapes of film cooling holes on the flow is not the same, requiring detailed numerical simulation and experimental verification; (2) Changing the arrangement of the film cooling holes will affect the continuity and integrity of the film cooling, and the degree of interference of different arrangement of film cooling holes on the flow is different, requiring reasonable optimization design and parameter selection; (3) Changing the inclination angle of the film cooling holes will affect the flow direction and velocity of the film cooling, and the adhesion and cooling effect of film cooling holes with different inclination angles are different, requiring precise control and adjustment.

[0007] Given these challenges, developing a novel gradient aperture film cooling (FSC) layout for turbine blade gill regions is particularly important. This structure needs to be able to adapt to the complex flow characteristics of the gill regions, and achieve precise control of the cold air outflow by adjusting the size and distribution of the FSC apertures, thereby improving the uniformity of the film coverage and the cooling effect. Summary of the Invention

[0008] (I) Purpose of the Invention

[0009] To address the aforementioned deficiencies and shortcomings of existing technologies, particularly the problem of uneven film cooling distribution in the gill region of aero-engine turbine blades due to complex flow structures, traditional uniform film cooling technology cannot provide sufficient thermal protection, thus affecting blade reliability and service life. This invention aims to propose a gradient aperture film cooling layout structure for the gill region of aero-engine turbine blades. By designing the film aperture in the gill region (the area adjacent to the leading edge with large curvature changes) according to the velocity distribution along the blade span, the film aperture gradually increases or decreases from the blade root to the blade tip in a certain gradient, achieving complete and good film coverage of the gill region. This gradient aperture layout not only considers the local characteristics of the flow but also enhances the cooling effect through precise control of the film outflow, improving the utilization efficiency and uniformity of the cooling gas, thereby significantly improving the thermal protection performance and high-temperature resistance of the turbine blades. Therefore, this invention has the advantages of strong functionality and clear application, which can effectively improve the film cooling performance of the turbine blade gill area, reduce the surface temperature and thermal stress of the blade, improve the reliability and life of the blade, and provide a new solution for high-efficiency and high-performance aero-engine turbine blade cooling technology.

[0010] (II) Technical Solution

[0011] To achieve the objective of this invention and solve its technical problems, the present invention adopts the following technical solution:

[0012] A gradient aperture film cooling layout structure for the gill region of an aero-engine turbine blade includes a plurality of turbine blades uniformly distributed circumferentially. Each turbine blade includes, in the chord direction, a gill region with a large curvature variation, distributed between the leading edge and the blade body. The structure is characterized by...

[0013] The blade surface of the gill region is provided with a plurality of air film pores arranged in an array along the blade span from the blade root to the blade tip. Based on the velocity distribution characteristics along the blade span on the blade surface of the gill region, the diameter of each air film pore in each row gradually increases or decreases along the blade span from the blade root to the blade tip according to a preset pore size gradient ε, either as a single air film pore or as a group of several adjacent air film pores in the span direction. This achieves complete and good air film coverage of the blade surface of the gill region and enhances the cooling effect. During the process of gradually increasing or decreasing the air film pore diameter according to the preset pore size gradient ε, the pore diameter of a single air film pore or a group of air film pores located at the blade root is used as the reference pore diameter D. Air film pores in the same group of air film pores have the same pore diameter. The pore diameter gradually increases or decreases from the blade root to the blade tip according to the following expression: D1, D2, ..., D... i :

[0014]

[0015] The aperture gradient ε is defined as ΔD / Δy, which is the ratio of the difference between adjacent apertures ΔD to the difference between their corresponding spanwise positions Δy. i and D i-1 These represent the aperture diameters of two adjacent film pores or groups of two adjacent film pores along the blade span, y i and y i-1 The reference aperture D is set between 0.6 mm and 2.0 mm, representing the spanwise position of two adjacent film vents or two adjacent film vent groups relative to the blade root in the blade spanwise direction. The specific value of the aperture gradient ε is optimized according to the heat load distribution and flow characteristics of the blade gill area to ensure that the jet of each film vent in the film vent row can form a stable and uniformly distributed film layer on the blade surface in the blade gill area, thereby achieving a high-efficiency cooling effect in this area.

[0016] Preferably, the aperture gradient ε is set based on the flow velocity distribution, heat load gradient, and blade surface temperature distribution in the blade gill region, and is set to a constant or variable value according to actual cooling requirements. This ensures that the increase or decrease in the film gas pore diameter matches the heat load distribution and fluid flow characteristics along the spanwise direction of the blade gill region, achieving uniform distribution and stable coverage of the film gas. During implementation, the specific value of the aperture gradient ε is determined through numerical simulation or experimental verification to ensure its cooling effect in practical applications.

[0017] Preferably, the air film pores on the leaf surface of the leaf gill region are arranged in a staggered or interlaced manner, and the center line of each air film pore is not at the same spanwise height position as the center line of the upper or lower air film pore, so as to increase the opening area of ​​the air film pores and the air film coverage.

[0018] Preferably, the spacing P between two adjacent film pores in the same film pore row in the blade spanwise direction is set between 3D and 5D according to the specific cooling requirements of the blade gill area, where D is the reference aperture of the film pore, so as to ensure that the jets of adjacent film pores have a certain overlap and interference to form a continuous film layer and enhance the overall cooling efficiency of the blade gill area.

[0019] Furthermore, adjacent film cooling pores employ a varying spacing P along the blade's span, with P gradually increasing or decreasing from the leaf root to the leaf tip, to accommodate the varying flow velocity characteristics along the span in the leaf gill region. This varied spacing arrangement helps optimize the interaction between film cooling pores, adjusting the coverage and density of the film cooling system to achieve more uniform and effective blade surface cooling.

[0020] Preferably, the inlet end of each of the gas film holes is connected to the hollow cavity of the turbine blade filled with cooling gas, and the outlet end is connected to the main gas flow channel. The centerline of each gas film hole is set with an inclination angle θ relative to the blade surface. The inclination angle θ is between 30° and 60° to optimize the interaction between the gas film injection and the main gas flow, and enhance the adhesion and cooling capacity of the gas film.

[0021] Furthermore, the tilt angle θ of the air film pores is optimized according to the flow direction and intensity at different locations in the blade gill region to adjust the flow direction and diffusion range of the air film, thereby improving the adhesion and cooling effect of the air film. By precisely controlling the tilt angle of the air film pores, it can be ensured that the air film forms a continuous and stable covering layer along the blade surface, effectively isolating the high-temperature airflow from the thermal effects on the blade.

[0022] Preferably, the outlet of the air film hole expands in a fan shape along the airflow direction to increase the effective area of ​​the air film hole outlet, thereby reducing the jet velocity and momentum of the air film hole, enhancing the jet stability and adhesion of the air film layer, and reducing the detachment of the air film.

[0023] Preferably, the shape of the air film orifice is cylindrical or irregularly shaped, wherein the irregularly shaped orifice is elliptical, fan-shaped, rhomboid, or other non-circular, to adapt to the specific hydrodynamic and thermal load characteristics of the blade gill region. Cylindrical orifices are widely used due to their mature manufacturing process and relatively simple processing, and can provide stable air film coverage. Irregularly shaped orifices, on the other hand, optimize the jet characteristics of the airflow and the adhesion of the air film by changing the geometry of the orifice, thereby enhancing the stability and coverage of the air film, especially in high curvature areas or areas prone to flow separation on the blade surface.

[0024] Preferably, the air film pores are surrounded by microgrooves or microprotrusions to enhance the interaction between the air film and the blade surface, thereby improving the stability and cooling efficiency of the air film. These microstructures can increase the turbulence on the blade surface and improve the adhesion of the air film, thus forming a more stable and uniform cooling air film layer on the blade surface, effectively improving the cooling performance and high-temperature resistance of the blade.

[0025] (III) Technical Effects

[0026] Compared with the prior art, the gradient aperture film cooling layout structure of the turbine blade gill region of the present invention has the following beneficial and significant technical effects:

[0027] (1) This invention achieves efficient and complete film cooling coverage in the gill region of turbine blades by employing a gradient aperture design. Compared with the traditional uniform aperture layout, the gradient aperture layout fully considers the velocity variation and heat load distribution along the blade span in the gill region, enabling each film cooling hole to provide the most suitable cooling effect according to local cooling needs. This targeted design greatly improves cooling efficiency and ensures the structural integrity and material properties of the turbine blades under high-temperature environments. In addition, the film cooling holes used in this invention are cylindrical or irregularly shaped holes, which have relatively simple manufacturing processes, do not require complex processing equipment and processes, reduce manufacturing costs and difficulties, and improve manufacturing efficiency and quality.

[0028] (2) The gradient aperture air film cooling layout structure of the turbine blade gill area of ​​the present invention is specifically designed for the gill area of ​​the turbine blade with large curvature. In response to the drastic flow changes in the gill area of ​​the blade with large curvature, a gradient aperture air film cooling layout is designed so that the outflow rate and outflow velocity of the air film holes are matched with the flow, achieving efficient and complete air film coverage, effectively isolating the direct impact of high temperature gas on the blade, and significantly reducing the surface temperature and thermal stress of the blade.

[0029] (3) The gradient aperture air film cooling layout of the turbine blade gill area of ​​the present invention can flexibly arrange the air film holes from the blade root to the blade tip according to the geometry and size of the blade gill area, without being restricted by the space of the blade leading edge and the blade body, making full use of the cooling space of the blade gill area, improving the utilization efficiency and cooling uniformity of the cooling gas, and meeting the air film cooling requirements of the blade gill area.

[0030] (4) The gradient aperture air film cooling layout of the present invention is implemented in the gill region of the blade. Different aperture air film holes are set according to geometric and flow characteristics, resulting in a good improvement in cooling effect. Through the gradient aperture design, a layered structure of the air film is realized, which makes the air film have a strong interference effect and anti-blow-off ability. At the same time, it increases the flow kinetic energy and flow direction change of the air film, improves the adhesion and cooling effect of the air film, and reduces the flow loss and manufacturing difficulty of the air film. Attached Figure Description

[0031] Figure 1 The diagram shows a schematic of the gradient aperture film cooling layout structure of the gill region of an aero-engine turbine blade according to the present invention.

[0032] Figure 2 The image shows a top view (xz plane) of the gradient aperture air film cooling layout in the gill region, where x is the chord direction of the blade and z is the thickness direction of the blade.

[0033] Figure 3 The diagram shows a cross-sectional view (yz section) of the gradient aperture film cooling layout in the gill region, where y is the blade spanwise and z is the blade thickness direction.

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

[0035] 1-Turbine blade, 2-Blade gill area, 3-Film gas hole, 4-Film gas hole inlet end, 5-Film gas hole outlet end, 6-Main gas flow, 7-Film gas hole centerline, 8-Main gas flow direction, 9-Blade spanwise. Detailed Implementation

[0036] 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.

[0037] Figure 1 The diagram shows a schematic of the gradient aperture film cooling layout structure for the gill region of an aero-engine turbine blade according to the present invention. Figure 1 As shown, the gradient aperture air film cooling layout structure of the turbine blade gill region of the present invention includes a plurality of turbine blades 1 uniformly distributed along the circumference, and each turbine blade 1 includes a blade gill region 2 with large curvature variation characteristics distributed between the blade leading edge and the blade body in the chord direction. A plurality of air film holes 3 are arranged in an array along the blade span from the blade root to the blade tip on the blade surface of the blade gill region 2. Based on the velocity distribution characteristics along the blade span on the blade surface of the blade gill region 2, the diameter of each air film hole 3 in each air film hole row gradually increases or decreases along the blade span from the blade root to the blade tip according to a preset aperture gradient ε, either as a single air film hole or as a group of several adjacent air film holes in the span direction. This achieves complete and good air film coverage of the blade surface of the blade gill region and enhances the cooling effect. During the process of gradually increasing or decreasing the air film hole diameter according to the preset aperture gradient ε, the aperture of a single air film hole or a group of air film holes located at the blade root is used as the reference aperture diameter D. Air film holes in the same air film hole group have the same air film hole diameter. The aperture gradient ε gradually increases or decreases from the blade root to the blade tip according to the following expression: D1, D2, ..., D... i :

[0038]

[0039] The aperture gradient ε is defined as ΔD / Δy, which is the ratio of the difference between adjacent apertures ΔD to the difference between their corresponding spanwise positions Δy. i and D i-1 These represent the aperture diameters of two adjacent film pores or groups of two adjacent film pores along the blade span, y i and y i-1 The reference aperture D is set between 0.6 mm and 2.0 mm, representing the spanwise position of two adjacent film vents or two adjacent film vent groups relative to the blade root in the blade spanwise direction. The specific value of the aperture gradient ε is optimized according to the heat load distribution and flow characteristics of the blade gill area to ensure that the jet of each film vent in the film vent row can form a stable and uniformly distributed film layer on the blade surface in the blade gill area, thereby achieving a high-efficiency cooling effect in this area.

[0040] from Figure 1 As can be seen, compared with the conventional uniform distribution of air film pores from leaf root to leaf tip, the gradient pores in the leaf gill region gradually increase or decrease from leaf root to leaf tip, and are distributed with a certain gradient. The gradient pores can be individual air film pores, gradually increasing or decreasing from leaf root to leaf tip; or they can be in groups of 3-5 air film pores, with each group gradually increasing or decreasing from leaf root to leaf tip.

[0041] Figure 2 The diagram shows a top view (xz plane) of the gradient aperture film cooling layout in the gill region. It can be seen that the reference film cooling hole is a cylindrical structure with a diameter of D, which ranges from 0.6 to 2.0 mm. The gradient aperture film cooling layout in the gill region generally includes 1 to 2 film cooling holes. The hole spacing is P, which ranges from 3D to 5D, to ensure that the jets from adjacent film cooling holes 3 have a certain overlap and interference to form a continuous film layer, enhancing the overall cooling efficiency of the blade gill region. Furthermore, adjacent film cooling holes 3 can use a varying spacing P along the blade span 9, with P gradually increasing or decreasing from the blade root to the blade tip to adapt to the varying flow velocity characteristics along the span in the gill region 2. This varying spacing layout helps optimize the interaction between the film cooling holes, adjust the coverage and density of the film cooling, and achieve more uniform and effective blade surface cooling.

[0042] In a preferred embodiment of this invention, the aperture gradient ε is set based on the flow velocity distribution, heat load gradient, and blade surface temperature distribution in the blade gill region, and is adjusted according to actual cooling requirements. It can be constant or vary according to flow conditions, ensuring that the increase in the diameter of the film gas pores matches the heat load distribution and fluid flow characteristics along the spanwise direction of the blade gill region, achieving uniform distribution and stable coverage of the film gas layer. During implementation, the specific value of the aperture gradient ε is determined through numerical simulation or experimental verification to ensure its cooling effect in practical applications. The film gas pores on the blade surface in the blade gill region are arranged in a staggered or interleaved manner. Between adjacent rows of film gas pores, the centerline of each film gas pore is not at the same spanwise height as the centerline of the row above or below it, thereby increasing the opening area of ​​the film gas pores and the coverage range of the film gas layer.

[0043] Figure 3 The diagram shows a cross-sectional view (yz section) of the gradient aperture film cooling layout in the gill region. The angle between the centerline 7 of the film cooling aperture and the blade surface is θ, which is between 30° and 60° to optimize the interaction between the film cooling injection and the main combustion flow, thereby enhancing the adhesion and cooling capacity of the film cooling system. The tilt angle θ of the film cooling aperture is optimized according to the flow direction and intensity at different locations in the gill region 2 of the blade to adjust the flow direction and diffusion range of the film cooling system, thus improving the adhesion and cooling effect of the film cooling system. By precisely controlling the tilt angle of the film cooling aperture, a continuous and stable covering layer of film cooling can be formed along the blade surface.

[0044] In a preferred embodiment of the present invention, the film cooling layout with gradient apertures in the gill region employs film cooling holes 3 that can be cylindrical or irregularly shaped. Irregularly shaped holes can be elliptical, fan-shaped, rhomboid, or other non-circular, to adapt to the specific hydrodynamic and thermal load characteristics of the blade gill region. Cylindrical holes are widely used due to their mature manufacturing process and relatively simple processing, providing stable film coverage. Irregularly shaped holes, on the other hand, optimize the jet characteristics of the airflow and the adhesion of the film by changing the geometry of the holes, thereby enhancing the stability and coverage of the film, especially in high curvature areas or areas prone to flow separation on the blade surface. Furthermore, the outlet of the film cooling hole expands in a fan shape along the airflow direction to increase the effective area of ​​the outlet, thereby reducing the jet velocity and momentum of the film cooling hole, enhancing the jet stability and adhesion of the film layer, and reducing film layer rupture and peeling.

[0045] In a preferred embodiment of the invention, microgrooves or microprotrusions are provided around the air film vents to enhance the interaction between the air film and the blade surface, thereby improving the stability and cooling efficiency of the air film. These microstructures can increase the turbulence on the blade surface and improve the adhesion of the air film, thus forming a more stable and uniform cooling air film layer on the blade surface, effectively improving the cooling performance and high-temperature resistance of the blade.

[0046] It can be seen that the gradient aperture film cooling layout of the turbine blade gill region of the present invention can bring several advantages: 1. The gradient aperture film cooling layout of the gill region has a simple structure and is easy to process; 2. The gradient aperture film cooling layout of the gill region is specially set according to its geometric and flow characteristics, and the cooling effect is obvious; 3. The gradient aperture film cooling layout of the gill region can be flexibly arranged from the blade root to the blade tip, without being limited by space; 4. The gradient aperture film cooling layout of the gill region can effectively resist the adverse effects of large curvature changes, and ultimately ensure good film coverage on the surface of the turbine blade.

[0047] 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 gradient aperture film cooling layout structure for the gill region of an aero-engine turbine blade, comprising a plurality of turbine blades uniformly distributed circumferentially, each turbine blade including, in the chord direction, a gill region with a large curvature variation distributed between the leading edge and the blade body, characterized in that, The blade surface of the gill region is provided with a plurality of air film pores arranged in an array along the blade span from the blade root to the blade tip. Based on the velocity distribution characteristics along the blade span on the blade surface of the gill region, the diameter of each air film pore in each row gradually increases or decreases along the blade span from the blade root to the blade tip according to a preset pore size gradient ε, either as a single air film pore or as a group of several adjacent air film pores in the span direction. This achieves complete and good air film coverage of the blade surface of the gill region and enhances the cooling effect. During the process of gradually increasing or decreasing the air film pore diameter according to the preset pore size gradient ε, the pore diameter of a single air film pore or a group of air film pores located at the blade root is used as the reference pore diameter D. Air film pores in the same group of air film pores have the same pore diameter. The pore diameter gradually increases or decreases from the blade root to the blade tip according to the following expression: D1, D2, ..., D... i : The aperture gradient ε is defined as ΔD / Δy, which is the ratio of the difference between adjacent apertures ΔD to the difference between their corresponding spanwise positions Δy. i and D i-1 These represent the aperture diameters of two adjacent film pores or groups of two adjacent film pores along the blade span, y i and y i-1 The reference aperture D is set between 0.6 mm and 2.0 mm, representing the spanwise position of two adjacent film vents or two adjacent film vent groups relative to the blade root in the blade spanwise direction. The specific value of the aperture gradient ε is optimized according to the heat load distribution and flow characteristics of the blade gill area to ensure that the jet of each film vent in the film vent row can form a stable and uniformly distributed film layer on the blade surface in the blade gill area, thereby achieving a high-efficiency cooling effect in this area.

2. The gradient aperture film cooling layout structure for the gill region of an aero-engine turbine blade according to claim 1, characterized in that, The aperture gradient ε is set based on the flow velocity distribution, heat load gradient, and blade surface temperature distribution in the blade gill region, and is set to a constant or variable value according to actual cooling requirements. This ensures that the increase or decrease of the air film aperture diameter matches the heat load distribution and fluid flow characteristics along the span of the blade gill region, thereby achieving uniform distribution and stable coverage of the air film.

3. The gradient aperture film cooling layout structure for the gill region of aero-engine turbine blades according to claim 1, characterized in that, The air film pores on the leaf surface of the leaf gill region are arranged in a staggered or interlaced manner. Between two adjacent rows of air film pores, the center line of each air film pore is not at the same spanwise height position as the center line of the upper or lower row of air film pores, so as to increase the opening area of ​​the air film pores and the coverage of the air film.

4. The gradient aperture film cooling layout structure for the gill region of aero-engine turbine blades according to claim 1, characterized in that, The spacing P between two adjacent film pores in the same film pore row in the blade span direction is set between 3D and 5D according to the specific cooling requirements of the blade gill area, where D is the reference aperture of the film pore, so as to ensure that the jets of adjacent film pores have a certain overlap and interference to form a continuous film layer and enhance the overall cooling efficiency of the blade gill area.

5. The gradient aperture film cooling layout structure for the gill region of an aero-engine turbine blade according to claim 4, characterized in that, The spacing P between two adjacent air film pores varies along the leaf span, gradually increasing or decreasing from the leaf root to the leaf tip, to adapt to the varying flow velocity characteristics of the leaf gill region along the span.

6. The gradient aperture film cooling layout structure for the gill region of an aero-engine turbine blade according to claim 1, characterized in that, The inlet end of each of the aforementioned film gas holes is connected to the hollow cavity of the turbine blade filled with cooling gas, and the outlet end is connected to the main gas flow channel. The centerline of each film gas hole is set with an inclination angle θ relative to the blade surface. The inclination angle θ is between 30° and 60° to optimize the interaction between film gas injection and main gas flow, and enhance the adhesion and cooling capacity of the film gas.

7. The gradient aperture film cooling layout structure for the gill region of an aero-engine turbine blade according to claim 6, characterized in that, The tilt angle θ of the air film pores is optimized according to the flow direction and intensity at different positions in the leaf gill region to adjust the flow direction and diffusion range of the air film, thereby improving the adhesion and cooling effect of the air film.

8. The gradient aperture film cooling layout structure for the gill region of an aero-engine turbine blade according to claim 1, characterized in that, The outlet of the air film orifice expands in a fan shape along the airflow direction to increase the effective area of ​​the air film orifice outlet, thereby reducing the jet velocity and momentum of the air film orifice, enhancing the jet stability and adhesion of the air film, and reducing the phenomenon of air film detachment.

9. The gradient aperture film cooling layout structure for the gill region of an aero-engine turbine blade according to claim 1, characterized in that, The air film pores are cylindrical or irregularly shaped, wherein the irregularly shaped pores are elliptical, fan-shaped, rhomboid, or other non-circular pores, to adapt to the specific hydrodynamic and thermal load characteristics of the leaf gill region.

10. The gradient aperture film cooling layout structure for the gill region of an aero-engine turbine blade according to claim 1, characterized in that, The air film pores are surrounded by microgrooves or microprotrusions to enhance the interaction between the air film and the blade surface, thereby improving the stability of the air film and cooling efficiency.

Images