CN118030202A - Gradient pore air film cooling layout structure for gill region of turbine blade of aeroengine - Google Patents
Gradient pore air film cooling layout structure for gill region of turbine blade of aeroengine Download PDFInfo
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Abstract
The invention relates to a gradient aperture air film cooling layout structure of a turbine blade gill region of an aeroengine, which aims at providing an improved cooling effect aiming at the cooling requirement of a turbine blade under high-temperature, high-speed and high-pressure working conditions, in particular to a key region with larger curvature change and severe flow change in the blade gill region. According to the structure, the gradient-changing air film pore diameter is arranged in the fin area of the blade along the spanwise direction of the blade, so that the size of the air film pore is gradually adjusted according to the change of the flow velocity and the heat load distribution, and the distribution and the coverage effect of cooling air flow are optimized. The structure has the advantages of simple structure, strong pertinence, flexible arrangement, obvious improvement of cooling effect and the like, improves the cooling efficiency and cooling uniformity of the gill area of the blade, effectively prolongs the service life of the turbine blade, and improves the overall performance and reliability of the engine. The invention also provides a plurality of optimized air film hole shapes, arrangement modes and inclination angles so as to further improve the coverage rate, stability, adhesiveness and cooling effect of the air film.
Description
Technical Field
The invention belongs to the technical field of cooling of turbine blades of an aeroengine, relates to design and optimization of a gas film cooling layout structure of a turbine blade gill region, and particularly relates to a gradient pore diameter gas film cooling layout structure of the turbine blade gill region of the aeroengine.
Background
Aero-engine turbine blades are one of the core components of an engine and are extremely harsh in their operating environment and subjected to high temperature, high pressure, high velocity airflow shocks, and therefore effective cooling techniques are required to preserve the structural integrity and material properties of the blades. The air film cooling technology is a common turbine blade cooling technology, and the principle is that a series of air film holes are arranged on the surface of a blade, and an air film covering the surface of the blade is formed by cold air sprayed out of the air film holes, so that the direct impact of hot air on the blade is isolated, and the surface temperature and thermal stress of the blade are reduced.
In the existing air film cooling scheme of the turbine blade of the aeroengine, the layout generally comprises the steps that a plurality of rows of dense air film holes are arranged on the front edges of the pressure surface and the suction surface of the turbine blade to realize spray cooling, 3-6 rows of air film holes which are uniformly distributed are arranged on the pressure surface, 2-4 rows of air film holes which are uniformly distributed are arranged on the suction surface, and the overall heat protection of the blade is realized. The advantages of this layout are that it is simple and easy to implement, and it can meet the cooling requirements of the turbine blade to a certain extent, and protect the blade from direct attack by high-temperature gas.
However, with the pursuit of high efficiency and high performance of aero-engines, the temperature of the turbine inlet is continuously increased, and higher requirements are put on the cooling technology of turbine blades, so that the design requirements are hardly met any more by adopting the conventional uniform film cooling technology. In particular, the complex appearance and geometric parameters of the turbine blade are very complex, so that the flow structure is extremely complex, and the uniformly arranged air film holes are extremely easy to cause the partial area of the surface of the turbine blade to be uncovered by the air film, thereby affecting the reliability and the service life of the blade due to ablation. Among these, the turbine blade gill region (gill region, the transition region between the blade leading edge and the blade body) is particularly interesting. The curvature of the gill area is changed greatly, the pressure gradient is large, the flow is changed severely along the flow direction, and the flow speed from the blade root to the blade tip along the spanwise direction also can be changed greatly, so that the cooling layout design of the area faces a great challenge, and therefore, the targeted fine air film cooling layout is necessary.
Currently, research into film cooling arrangements for the gill region of turbine blades has focused mainly on the following aspects: (1) Changing the shape of the air film hole, such as adopting elliptical, fan-shaped, diamond-shaped and other non-circular air film holes, so as to increase the outlet area of the air film hole and improve the coverage rate and stability of the air film; (2) Changing the arrangement mode of the air film holes, such as adopting staggered arrangement, spiral arrangement and other non-uniformly arranged air film holes, so as to increase the interference effect of the air film and improve the blow-off resistance of the air film; (3) And the inclination angle of the air film holes is changed, such as different air films Kong Qingjiao with forward inclination, reverse inclination, compound inclination and the like are adopted, so that the flow direction and the flow speed of the air film are adjusted, and the adhesiveness and the cooling effect of the air film are improved.
Although the above studies have improved the film cooling performance of the gill region of the turbine blade to some extent, the following technical problems or challenges remain: (1) Changing the shape of the air film hole can increase the manufacturing difficulty and cost of the air film hole, and the influence of the air film holes with different shapes on the flow is different, so that careful numerical simulation and experimental verification are required; (2) Changing the arrangement mode of the air film holes can influence the continuity and the integrity of the air film, and the air film holes with different arrangement modes have different interference degrees on the flow, so that reasonable optimization design and parameter selection are required; (3) Changing the inclination of air film hole can influence the flow direction and the velocity of flow of air film, and the air film hole at different inclination is different to adhesion and the cooling effect of flow, needs to carry out accurate control and regulation.
In view of these challenges, it is particularly important to develop a novel turbine blade gill region gradient aperture air film cooling layout structure. The structure needs to be capable of adapting to the complex flow characteristics of the gill area, and fine control of cold air outflow is realized by adjusting the size and distribution of the air film holes so as to improve the covering uniformity and cooling effect of the air film.
Disclosure of Invention
Object of the invention
Aiming at the defects and the shortcomings in the prior art, particularly the problem that the distribution of air film outflow is uneven in a gill area of a turbine blade of an aeroengine due to a complex flow structure, the traditional uniform air film cooling technology cannot provide enough heat protection, and further the technical problem that the reliability and the service life of the blade are affected. Therefore, the invention has the advantages of strong functionality and definite purpose, can effectively improve the air film cooling performance of the gill area of the turbine blade, reduce the surface temperature and thermal stress of the blade, improve the reliability and service life of the blade, and provides a new solution for the high-efficiency and high-performance aeroengine turbine blade cooling technology.
(II) technical scheme
In order to achieve the aim of the invention and solve the technical problems, the invention adopts the following technical scheme:
A gradient pore air film cooling layout structure of an aeroengine turbine blade gill region comprises a plurality of turbine blades uniformly distributed along the circumferential direction, each turbine blade comprises a blade gill region with large curvature change characteristics distributed between the front edge of the blade and a blade main body in the chord direction,
According to the distribution characteristic of the flow velocity on the blade surface of the blade gill area along the blade span direction, the air film holes in each air film hole row have the same air film hole diameter, according to the single air film hole or the mode that a plurality of air film holes adjacent in the span direction are formed into an air film hole group, the air film hole diameter gradually increases or gradually decreases from the blade root to the blade tip along the blade span direction according to a preset aperture gradient epsilon, so that the purpose of covering the complete good air film of the blade surface of the blade gill area and achieving the enhanced cooling effect is achieved, wherein the aperture of the single air film hole or the air film hole group at the blade root is taken as a standard aperture D in the process of gradually increasing or gradually decreasing the air film hole diameter according to the preset aperture gradient epsilon, the air film hole diameter in the same air film hole group is gradually increased or gradually decreased from the blade root to the blade tip as D 1,D2,……,Di according to the aperture gradient epsilon shown in the following expression:
The aperture gradient epsilon is defined as delta D/delta y, namely the ratio of the difference delta D between adjacent apertures to the difference delta y between corresponding spanwise positions, D i and D i-1 are the apertures of two adjacent air film holes or two adjacent air film hole groups along the spanwise direction of the blade, y i and y i-1 are the spanwise positions of two adjacent air film holes or two adjacent air film hole groups relative to the blade root in the spanwise direction of the blade, the range of the reference aperture D is between 0.6mm and 2.0mm, the specific numerical value of the aperture gradient epsilon is optimally set according to the heat load distribution and the flow characteristics of the blade gill region, so that the jet flow of each air film hole in the air film hole row can form a stable and uniformly distributed air film layer on the surface of the blade gill region, and further the efficient cooling effect is realized in the region.
Preferably, the pore diameter gradient epsilon is set based on the flow velocity distribution, the heat load gradient and the blade surface temperature distribution of the blade gill region and is set to a constant or variable value according to actual cooling requirements, so that the increment or decrement of the diameter of the air film pore matches the heat load distribution and the fluid flow characteristic of the blade gill region along the spanwise direction, and uniform distribution and stable coverage of the air film are realized. In practice, the specific value of the pore size gradient epsilon is determined by numerical simulation or experimental verification to ensure the cooling effect in practical application.
Preferably, the air film holes on the blade surface of the blade gill region are staggered or staggered, and the central line of each air film hole is not positioned at the same spanwise height position with the central line of the upper or lower exhaust film hole between two adjacent air film holes, so as to increase the open area of the air film holes and the air film coverage range.
Preferably, the distance P between two adjacent air film holes in the same air film hole row in the blade expanding direction is set between 3D and 5D according to the specific cooling requirement of the blade gill area, wherein D is the reference aperture of the air film holes, so that certain overlapping and interference between jet flows of the adjacent air film holes are ensured to form a continuous air film layer, and the overall cooling efficiency of the blade gill area is enhanced.
Further, the adjacent two air film holes adopt a variable interval P in the spanwise direction of the blade, and the interval P gradually increases or decreases from the blade root to the blade tip so as to adapt to the variable characteristic of the flow velocity of the gill region of the blade along the spanwise direction. This varied pitch arrangement helps to optimize interactions between film holes, adjust film coverage and density, and achieve more uniform and efficient blade surface cooling.
Preferably, the inlet end of each air film hole is communicated with the hollow cavity filled with cooling gas of the turbine blade, the outlet end of each air film hole is communicated with the main gas flow channel, and the central line of each air film hole is provided with an inclined angle theta relative to the surface of the blade, and the inclined angle theta ranges from 30 degrees to 60 degrees so as to optimize the interaction between air film injection and the main gas flow and enhance the adhesiveness and cooling capability of the air film.
Further, the inclination angle theta of the air film hole is optimally set according to the flow direction and the strength of different positions of the fin area of the blade so as to adjust the flow direction and the diffusion range of the air film, thereby improving the adhesiveness and the cooling effect of the air film. Through the inclination of accurate control air film hole, can ensure that the air film forms continuous stable overburden along the blade surface, effectively the thermal influence of isolated high temperature air current to the blade.
Preferably, the outlet of the air film hole is fanned and expanded along the air flow direction to increase the effective area of the outlet of the air film hole, so as to reduce the jet speed and momentum of the air film hole, enhance the jet stability of the air film hole and the adhesiveness of the air film layer, and reduce the detachment phenomenon of the air film.
Preferably, the shape of the air film hole is a cylindrical hole or a special-shaped hole, wherein the special-shaped hole is an elliptical hole, a fan-shaped hole, a diamond-shaped hole or other non-circular holes so as to adapt to specific fluid dynamics and heat load characteristics of the gill region of the blade. The cylindrical holes are widely used because of their mature manufacturing process and relatively simple processing, and can provide stable air film coverage. The special-shaped holes optimize the jet characteristic of the air flow and the adhesive capability of the air film by changing the geometric shape of the holes, so that the stability and coverage of the air film are enhanced, and particularly, the air film is in a high-curvature area or an area where flow separation is easy to occur on the surface of the blade.
Preferably, micro grooves or micro protrusions are arranged around the air film holes so as to enhance the interaction between the air film and the surface of the blade and improve the stability and cooling efficiency of the air film. The microstructures can increase the turbulence degree of the surface of the blade and improve the adhesiveness of the air film, so that a more stable and uniform cooling air film layer is formed on the surface of the blade, and the cooling performance and the high temperature resistance of the blade are effectively improved.
(III) technical effects
Compared with the prior art, the gradient pore air film cooling layout structure for the gill region of the turbine blade of the aeroengine has the following beneficial and remarkable technical effects:
(1) According to the invention, by adopting a gradient aperture design, efficient and complete air film coverage is realized in the gill region of the turbine blade. Compared with the traditional uniform aperture layout, the gradient aperture layout fully considers the flow velocity change and the heat load distribution of the fin region of the blade along the span direction of the blade, so that each air film hole can provide an optimal cooling effect according to local cooling requirements. The targeted design greatly improves the cooling efficiency and ensures the structural integrity and the material performance of the turbine blade in a high-temperature environment. In addition, the air film hole adopted by the invention is a cylindrical hole or a special-shaped hole, the manufacturing process is relatively simple, complex processing equipment and process are not needed, the manufacturing cost and difficulty are reduced, and the manufacturing efficiency and quality are improved.
(2) The gradient pore air film cooling layout structure of the turbine blade gill region of the aeroengine is specially used for the blade gill region with larger curvature of the turbine blade, and the air film cooling layout of the gradient pore is designed aiming at severe flow change of the blade gill region with large curvature change characteristics, so that the output flow and the output flow speed of air film holes are matched with the flow, high-efficiency complete air film coverage is realized, direct impact of high-temperature gas on the blade is effectively isolated, and the surface temperature and the thermal stress of the blade are remarkably reduced.
(3) According to the gradient pore air film cooling layout of the blade gill region of the aeroengine turbine, air film holes can be flexibly arranged from blade roots to blade tips according to the geometric shape and the size of the blade gill region, the air film cooling layout is not limited by the space of the front edge of the blade and the space of the blade main body, the cooling space of the blade gill region is fully utilized, the utilization efficiency and the cooling uniformity of cooling air are improved, and the air film cooling requirement of the blade gill region is met.
(4) The gradient pore diameter air film cooling layout is intensively implemented in the gill region of the blade, and air film holes with different pore diameters are arranged according to geometric and flow characteristics, so that the cooling effect is improved well. Through the design of gradient aperture, realized the layered structure of air film for the air film has stronger interference effect and anti-blow-off ability, has increased the flow kinetic energy and the flow direction change of air film simultaneously, has improved the adhesion and the cooling effect of air film, has reduced the flow loss and the manufacturing degree of difficulty of air film simultaneously.
Drawings
FIG. 1 is a schematic diagram of a gradient pore air film cooling layout structure in a gill region of a turbine blade of an aircraft engine.
Fig. 2 shows a top view (x-z plane) of the gill zone gradient pore-size film cooling layout, where x is the blade chord direction and z is the blade thickness direction.
Fig. 3 shows a sectional view of a gill zone gradient pore-size film cooling layout (y-z section), where y is the spanwise direction of the blade and z is the thickness direction of the blade.
Reference numerals illustrate:
The turbine blade comprises a turbine blade body, a blade gill area 2, a gas film hole 3, a gas film Kong Jinkou end 4, a gas film hole outlet end 5, a main gas flow 6, a gas film hole center line 7, a main gas flow 8 and a blade direction-expanding 9.
Detailed Description
For a better understanding of the present invention, the following examples are set forth to illustrate the present invention. In the drawings, the same or similar reference numerals denote the same or similar elements or elements having the same or similar functions throughout. The described embodiments are some, but not all, embodiments of the invention. The embodiments described below by referring to the drawings are illustrative and intended to explain the present invention and should not be construed as limiting the invention. All other embodiments, which can be made by those skilled in the art based on the embodiments of the invention without making any inventive effort, are intended to be within the scope of the invention. The following describes the structure and technical scheme of the present invention in detail with reference to the accompanying drawings, and an embodiment of the present invention is given.
FIG. 1 is a schematic diagram of a gradient pore air film cooling layout structure in a gill region of a turbine blade of an aircraft engine. As shown in fig. 1, the gradient pore air film cooling layout structure of the turbine blade gill region of the aeroengine comprises a plurality of turbine blades 1 which are uniformly distributed along the circumferential direction, wherein each turbine blade 1 comprises a blade gill region 2 with large curvature change characteristics which is distributed and arranged between the front edge of the blade and the blade main body in the chord direction. According to the distribution characteristic of the flow velocity on the blade surface of the blade gill region 2 along the blade span direction, the air film holes 3 in each air film hole row have the same air film hole diameter, according to the single air film hole or the mode that a plurality of air film holes adjacent in the span direction are formed into an air film hole group, the air film hole diameter gradually increases or gradually decreases from the blade root to the blade tip along the blade span direction according to a preset aperture gradient epsilon to achieve the purpose of covering the complete good air film on the blade surface of the blade gill region and achieving the enhanced cooling effect, wherein the aperture of the single air film hole or the air film hole group at the blade root is taken as a reference aperture D in the process of gradually increasing or gradually decreasing the air film hole diameter according to the preset aperture gradient epsilon, and the air film hole diameter in the same air film hole group gradually increases or gradually decreases from the blade root to the blade tip to D 1,D2,……,Di as shown in the following expression:
The aperture gradient epsilon is defined as delta D/delta y, namely the ratio of the difference delta D between adjacent apertures to the difference delta y between corresponding spanwise positions, D i and D i-1 are the apertures of two adjacent air film holes or two adjacent air film hole groups along the spanwise direction of the blade, y i and y i-1 are the spanwise positions of two adjacent air film holes or two adjacent air film hole groups relative to the blade root in the spanwise direction of the blade, the range of the reference aperture D is between 0.6mm and 2.0mm, the specific numerical value of the aperture gradient epsilon is optimally set according to the heat load distribution and the flow characteristics of the blade gill region, so that the jet flow of each air film hole in the air film hole row can form a stable and uniformly distributed air film layer on the surface of the blade gill region, and further the efficient cooling effect is realized in the region.
As can be seen from fig. 1, compared with the conventional uniform arrangement of the air film holes from the blade root to the blade tip, the gradient holes of the blade gill region 2 gradually increase or decrease from the blade root to the blade tip and are distributed with a certain gradient change. The gradient holes can be gradually increased or gradually decreased from the blade root to the blade tip by taking a single air film hole as a unit; and 3-5 air film holes can be formed into a group, and each group gradually increases or gradually decreases from the blade root to the blade tip.
FIG. 2 shows a top view (x-z plane) of a gill zone gradient pore-size film cooling layout. It can be seen that the reference air film hole is of a cylindrical structure, the diameter is D, and D is between 0.6 and 2.0 mm. The gill zone gradient pore size air film cooling arrangement generally comprises 1-2 exhaust air film holes. The hole spacing is P, and the P is between 3D and 5D, so that certain overlapping and interference among jet flows of adjacent air film holes 3 are ensured to form a continuous air film layer, and the overall cooling efficiency of the fin area of the blade is enhanced. In addition, the adjacent two air film holes 3 can adopt a variable interval P on the spanwise direction 9 of the blade, and the interval P gradually increases or decreases from the blade root to the blade tip so as to adapt to the variable characteristic of the flow velocity of the gill region 2 of the blade along the spanwise direction. This varied pitch arrangement helps to optimize interactions between film holes, adjust film coverage and density, and achieve more uniform and efficient blade surface cooling.
In the preferred embodiment of the invention, the pore diameter gradient epsilon is set based on the flow velocity distribution, the heat load gradient and the blade surface temperature distribution of the blade gill region and can be set according to actual cooling requirements, and can be constant or can be changed according to the flow condition, so that the increment of the diameter of the air film pore is matched with the heat load distribution and the fluid flow characteristic of the blade gill region along the expanding direction, and the uniform distribution and the stable coverage of the air film layer are realized. In practice, the specific value of the pore size gradient epsilon is determined by numerical simulation or experimental verification to ensure the cooling effect in practical application. The arrangement mode of the air film holes on the blade surface of the blade gill area is staggered or staggered, and the central line of each air film hole is not positioned at the same spanwise height position with the central line of the upper row or the lower exhaust film hole between two adjacent rows of air film holes so as to increase the open pore area of the air film holes and the air film coverage range.
Fig. 3 shows a sectional view (y-z section) of a gill region gradient aperture air film cooling layout, wherein an included angle between the center line 7 of an air film hole and the surface of a blade is theta, and theta is between 30 and 60 degrees, so that the interaction between air film injection and main gas flow is optimized, and the adhesiveness and cooling capacity of the air film are enhanced. The inclination angle theta of the air film holes is optimally set according to the flow directions and the intensities of different positions of the blade gill region 2 so as to adjust the flow direction and the diffusion range of the air film, thereby improving the adhesiveness and the cooling effect of the air film. By precisely controlling the inclination angle of the air film holes, the air film can be ensured to form a continuous and stable covering layer along the surface of the blade.
In the preferred embodiment of the invention, the air film holes 3 used in the gill region gradient pore air film cooling layout can be cylindrical holes or special-shaped holes, wherein the special-shaped holes are elliptical holes, fan-shaped holes, diamond-shaped holes or other non-circular holes so as to adapt to specific fluid dynamics and thermal load characteristics of the gill region of the blade. The cylindrical holes are widely used because of their mature manufacturing process and relatively simple processing, and can provide stable air film coverage. The special-shaped holes optimize the jet characteristic of the air flow and the adhesive capability of the air film by changing the geometric shape of the holes, so that the stability and coverage of the air film are enhanced, and particularly, the air film is in a high-curvature area or an area where flow separation is easy to occur on the surface of the blade. In addition, the outlet of the air film hole is expanded in a fan shape along the air flow direction to increase the effective area of the outlet of the air film hole, so as to reduce the jet flow speed and momentum of the air film hole, enhance the jet flow stability of the air film hole and the adhesiveness of the air film layer, and reduce the cracking and stripping phenomena of the air film layer.
In the preferred embodiment of the invention, the circumference of the air film hole is provided with micro grooves or micro bulges so as to enhance the interaction between the air film and the surface of the blade and improve the stability and cooling efficiency of the air film. The microstructures can increase the turbulence degree of the surface of the blade and improve the adhesiveness of the air film, so that a more stable and uniform cooling air film layer is formed on the surface of the blade, and the cooling performance and the high temperature resistance of the blade are effectively improved.
It can be seen that the adoption of the gradient pore air film cooling layout of the gill area of the turbine blade of the aeroengine can bring about several benefits: 1. the gill region gradient pore diameter air film cooling layout structure is simple and convenient to process; 2. the cooling layout of the gill region gradient pore-diameter air film is specially set according to the geometric characteristics and the flow characteristics, and the cooling effect is obvious; 3. the gill area gradient pore diameter air film cooling layout is flexibly arranged along the blade root to the blade tip, and is not limited by space; 4. the cooling layout of the gradient pore-size air film in the gill area can effectively resist the adverse effect caused by large curvature change, and finally ensures that the air film on the surface of the turbine blade is well covered.
The object of the present invention is fully effectively achieved by the above-described embodiments. Those skilled in the art will appreciate that the present invention includes, but is not limited to, those illustrated in the drawings and described in the foregoing detailed description. While the invention has been described in connection with what is presently considered to be the most practical and preferred embodiment, it is to be understood that the invention is not limited to the disclosed embodiment, but on the contrary, is intended to cover various modifications and equivalent arrangements included within the scope of the appended claims.
Claims (10)
1. A gradient pore air film cooling layout structure of an aeroengine turbine blade gill region comprises a plurality of turbine blades uniformly distributed along the circumferential direction, each turbine blade comprises a blade gill region with large curvature change characteristics distributed between the front edge of the blade and a blade main body in the chord direction,
According to the distribution characteristic of the flow velocity on the blade surface of the blade gill area along the blade span direction, the air film holes in each air film hole row have the same air film hole diameter, according to the single air film hole or the mode that a plurality of air film holes adjacent in the span direction are formed into an air film hole group, the air film hole diameter gradually increases or gradually decreases from the blade root to the blade tip along the blade span direction according to a preset aperture gradient epsilon, so that the purpose of covering the complete good air film of the blade surface of the blade gill area and achieving the enhanced cooling effect is achieved, wherein the aperture of the single air film hole or the air film hole group at the blade root is taken as a reference aperture D in the process of gradually increasing or gradually decreasing the air film hole diameter according to the preset aperture gradient epsilon, the air film hole diameter in the same air film hole group is gradually increased or decreased from the blade root to the blade tip according to the aperture gradient epsilon shown in the following expression:
The aperture gradient epsilon is defined as delta D/delta y, namely the ratio of the difference delta D between adjacent apertures to the difference delta y between corresponding spanwise positions, D i and D i-1 are the apertures of two adjacent air film holes or two adjacent air film hole groups along the spanwise direction of the blade, y i and y i-1 are the spanwise positions of two adjacent air film holes or two adjacent air film hole groups relative to the blade root in the spanwise direction of the blade, the range of the reference aperture D is between 0.6mm and 2.0mm, the specific numerical value of the aperture gradient epsilon is optimally set according to the heat load distribution and the flow characteristics of the blade gill region, so that the jet flow of each air film hole in the air film hole row can form a stable and uniformly distributed air film layer on the surface of the blade gill region, and further the efficient cooling effect is realized in the region.
2. The air film cooling layout structure according to claim 1, wherein the pore size gradient epsilon is set based on the flow velocity distribution, the heat load gradient and the blade surface temperature distribution of the blade gill region and is set to a constant or variable value according to actual cooling requirements, so that the increment or decrement of the diameter of the air film pore matches the heat load distribution and the fluid flow characteristics of the blade gill region along the expanding direction, and uniform distribution and stable coverage of the air film are realized.
3. The air film cooling layout structure of gradient pore diameter in the gill region of turbine blades of an aeroengine according to claim 1, wherein the air film holes on the surface of the blades of the gill region of the blades are staggered or staggered, and the center line of each air film hole is not at the same spanwise height position with the center line of the upper or lower exhaust film hole between two adjacent rows of air film holes, so as to increase the open area and the air film coverage of the air film holes.
4. The air film cooling layout structure of the gradient pore diameter of the air film in the fin area of the turbine blade of the aeroengine according to claim 1, wherein the distance P between two adjacent air film holes in the same air film hole row in the fin span direction is set between 3D and 5D according to the specific cooling requirement of the fin area of the blade, wherein D is the reference pore diameter of the air film holes, so that certain overlapping and interference between jet flows of the adjacent air film holes are ensured to form a continuous air film layer, and the overall cooling efficiency of the fin area of the blade is enhanced.
5. The air engine turbine blade gill region gradient aperture air film cooling layout structure according to claim 4, wherein the adjacent two air film holes adopt a variable distance P in the spanwise direction of the blade, and the distance P gradually increases or decreases from the blade root to the blade tip so as to adapt to the variable characteristic of the flow velocity of the blade gill region along the spanwise direction.
6. The aircraft engine turbine blade gill zone gradient pore size air film cooling arrangement of claim 1, wherein the inlet end of each air film hole is communicated with the hollow cavity of the turbine blade filled with cooling air, the outlet end is communicated with the main gas flow channel, and the center line of each air film hole is provided with an inclination angle θ with respect to the blade surface, the inclination angle θ ranges from 30 ° to 60 °, so as to optimize the interaction of air film injection and main gas flow, and enhance the adhesiveness and cooling capability of the air film.
7. The air engine turbine blade gill region gradient aperture air film cooling layout structure according to claim 6, wherein the inclination angle θ of the air film holes is optimally set according to the flow direction and the strength of different positions of the blade gill region, so as to adjust the flow direction and the diffusion range of the air film, and further improve the adhesiveness and the cooling effect of the air film.
8. The air film cooling layout structure for the gradient pore diameter of the gill region of the turbine blade of the aeroengine according to claim 1, wherein the outlet of the air film hole is fanned along the air flow direction to increase the effective area of the outlet of the air film hole, so as to reduce the jet flow speed and momentum of the air film hole, enhance the jet flow stability of the air film hole and the adhesiveness of the air film, and reduce the detachment phenomenon of the air film.
9. The aircraft engine turbine blade gill region gradient pore air film cooling layout structure of claim 1, wherein the air film holes are cylindrical holes or profiled holes, wherein the profiled holes are elliptical holes, fan-shaped holes, diamond-shaped holes or other non-circular holes to adapt to specific hydrodynamic and thermal load characteristics of the blade gill region.
10. The aircraft engine turbine blade gill region gradient aperture air film cooling layout structure according to claim 1, wherein micro grooves or micro protrusions are arranged around the air film holes so as to enhance the interaction between the air film and the blade surface and improve the stability and cooling efficiency of the air film.
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| CN202410286816.9A CN118030202A (en) | 2024-03-13 | 2024-03-13 | Gradient pore air film cooling layout structure for gill region of turbine blade of aeroengine |
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| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202410286816.9A CN118030202A (en) | 2024-03-13 | 2024-03-13 | Gradient pore air film cooling layout structure for gill region of turbine blade of aeroengine |
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