CN115492643A - Aeroengine turbine guide blade flange plate cooling structure - Google Patents
Aeroengine turbine guide blade flange plate cooling structure Download PDFInfo
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- CN115492643A CN115492643A CN202211421825.1A CN202211421825A CN115492643A CN 115492643 A CN115492643 A CN 115492643A CN 202211421825 A CN202211421825 A CN 202211421825A CN 115492643 A CN115492643 A CN 115492643A
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- Prior art keywords
- holes
- cooling structure
- turbine guide
- cavity
- air film
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Classifications
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D9/00—Stators
- F01D9/02—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D25/00—Component parts, details, or accessories, not provided for in, or of interest apart from, other groups
- F01D25/08—Cooling; Heating; Heat-insulation
- F01D25/12—Cooling
-
- Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
- Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
- Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
- Y02T50/00—Aeronautics or air transport
- Y02T50/60—Efficient propulsion technologies, e.g. for aircraft
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- Engineering & Computer Science (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Turbine Rotor Nozzle Sealing (AREA)
Abstract
The application belongs to the technical field of non-variable-volume engine design, concretely relates to aeroengine turbine guide vane flange plate cooling structure, include: a flange having a cavity therein, an outer sidewall having a plurality of rows of impingement holes, and an inner sidewall having a plurality of rows of film holes; the multi-drain flow column is arranged in the cavity and supported between the outer side wall and the inner side wall; wherein, each row of the impact holes are communicated with the cavity and distributed in a staggered way; each exhaust film hole is communicated with the cavity and distributed in a staggered way; the drain flow columns are distributed in a staggered manner; each impact hole is surrounded by four turbulence columns; each air film hole is surrounded by four turbulence columns; a turbulent flow column is arranged between the adjacent impact holes and the air film holes.
Description
Technical Field
The application belongs to the technical field of non-variable-volume engine design, and particularly relates to a cooling structure for a guide vane edge plate of a turbine of an aircraft engine.
Background
Turbine guide vane's point portion among the aeroengine, the flange has been cup jointed to the root, the flange bears higher temperature load in the aeroengine working process, currently, do not receive high temperature damage for the protection flange, set up a plurality of gas film holes on the flange, let in the cooling gas in to the flange through the gas film hole, cool off it and form the gas film at flange inside wall face, with this protection flange not receive high temperature damage, as shown in fig. 1, this kind of technical scheme has following defect:
1) The cooling efficiency is low, and a large amount of cooling gas needs to be consumed;
2) All the air film holes are uniformly distributed, the edge plate is adaptive to the blade profile, the whole edge plate is in a diamond shape, the diagonal area is difficult to be effectively cooled, and a local high-temperature area is easy to form, so that the edge plate is damaged by high temperature.
The present application has been made in view of the above-mentioned technical drawbacks.
It should be noted that the above background disclosure is only used for assisting the understanding of the inventive concept and technical solutions of the present invention, and does not necessarily belong to the prior art of the present application, and the background disclosure should not be used for evaluating the novelty and creativity of the present application in the case that there is no clear evidence that the above content is disclosed at the filing date of the present application.
Disclosure of Invention
It is an object of the present application to provide an aeroengine turbine guide vane platform cooling arrangement that overcomes or mitigates at least one aspect of the technical disadvantages known to exist.
The technical scheme of the application is as follows:
an aeroengine turbine guide vane platform cooling structure comprising:
a flange plate having a cavity therein, a plurality of rows of impingement holes in an outer sidewall thereof, and a plurality of rows of air film holes in an inner sidewall thereof;
the multi-row turbulence column is arranged in the cavity and supported between the outer side wall and the inner side wall;
wherein the content of the first and second substances,
each row of the impact holes are communicated with the cavity and distributed in a staggered way;
each exhaust film hole is communicated with the cavity and distributed in a staggered way;
each drain flow column is distributed in a staggered way;
each impact hole is surrounded by four turbulence columns;
each air film hole is surrounded by four turbulence columns;
a turbulent flow column is arranged between the adjacent impact holes and the air film holes.
According to at least one embodiment of the application, in the aeroengine turbine guide vane platform cooling structure, the diameter of the impingement holes is 1.75-2.75 times of that of the film holes;
the diameter of the turbulent flow column is 2.5 times of the pore diameter of the air film pore.
According to at least one embodiment of the application, in the aeroengine turbine guide vane platform cooling structure, the distance between the turbulence columns is 6-7.5 times of the pore diameter of the film hole.
According to at least one embodiment of the application, in the aeroengine turbine guide vane platform cooling structure, the film holes in the diagonal area of the platform have high density distribution.
According to at least one embodiment of the application, in the aeroengine turbine guide vane platform cooling structure, the distance between the cavity and the edge of the platform is 3 times of the diameter of the film hole.
According to at least one embodiment of the application, in the cooling structure of the aeroengine turbine guide vane platform, each exhaust film hole is inclined towards the rear end of the platform.
According to at least one embodiment of the application, in the aeroengine turbine guide vane platform cooling structure, the outlet of each exhaust film hole is expanded towards the rear end of the platform.
According to at least one embodiment of the application, in the aeroengine turbine guide vane platform cooling structure, the diameter of each turbulence column gradually shrinks from the middle part to the two ends.
The application has at least the following beneficial technical effects:
the utility model provides an aeroengine turbine guide blade flange cooling structure, when aeroengine during operation, each impact hole of accessible lets in the cooling gas in to the cavity, the cooling gas that gets into the cavity can carry out impingement cooling to the flange inside wall, it flows to form the torrent, form the opposite vortex system in the middle of the vortex post that corresponds, form and revolve to opposite passageway whirlpool, with this can carry out abundant torrent heat transfer in the cavity, and can circle round and strike flange lateral wall, and carry out convection current disturbance heat transfer between the lateral wall, the effective heat transfer area of cooling gas in the cavity has been increased, final cooling gas is discharged through each exhaust film hole, form the gas film on the inboard surface of flange, the protection flange does not receive high temperature damage, higher cooling efficiency has to the flange, can greatly reduce the consumption to the cooling gas.
Drawings
FIG. 1 is a schematic illustration of a prior art aero-engine turbine guide vane platform cooling configuration;
FIG. 2 is a schematic illustration of an aircraft engine turbine guide vane platform cooling configuration provided by an embodiment of the present application;
FIG. 3 is a sectional view taken along line C-C of FIG. 2;
wherein:
1-a flange;
2-a turbulence column;
3-impact holes;
4-air film hole;
5-short diagonal region;
6-long diagonal region.
For better illustration of the present embodiment, certain elements of the drawings may be omitted, enlarged or reduced, and do not represent the size of an actual product, and furthermore, the drawings are for illustrative purposes only and should not be construed as limiting the present application.
Detailed Description
In order to make the technical solutions and advantages of the present application clearer, the technical solutions of the present application will be further clearly and completely described in the following detailed description with reference to the accompanying drawings, and it should be understood that the specific embodiments described herein are only some of the embodiments of the present application, and are only used for explaining the present application, but not limiting the present application. It should be noted that, for convenience of description, only the parts related to the present application are shown in the drawings, other related parts may refer to general designs, and the embodiments and technical features in the embodiments in the present application may be combined with each other to obtain a new embodiment without conflict.
In addition, unless otherwise defined, technical or scientific terms used in the description of the present application shall have the ordinary meaning as understood by one of ordinary skill in the art to which the present application belongs. The terms "upper", "lower", "left", "right", "center", "vertical", "horizontal", "inner", "outer", and the like used in the description of the present application, which indicate orientations, are used only to indicate relative directions or positional relationships, and do not imply that the devices or elements must have a specific orientation, be constructed and operated in a specific orientation, and when the absolute position of the object to be described is changed, the relative positional relationships may be changed accordingly, and thus, should not be construed as limiting the present application. The use of "first," "second," "third," and the like in the description of the present application is for descriptive purposes only to distinguish between different components and is not to be construed as indicating or implying relative importance. The use of the terms "a," "an," or "the" and similar referents in the description of the application should not be construed as an absolute limitation of quantity, but rather as the presence of at least one. The word "comprising" or "comprises", and the like, when used in this description, is intended to specify the presence of stated elements or items, but not the exclusion of other elements or items.
Further, it is noted that, unless expressly stated or limited otherwise, the terms "mounted," "connected," and the like are used in the description of the invention in a generic sense, e.g., connected as either a fixed connection or a removable connection or integrally connected; can be mechanically or electrically connected; they may be directly connected or indirectly connected through an intermediate medium, or they may be connected through the inside of two elements, and those skilled in the art can understand their specific meaning in this application according to the specific situation.
The present application is described in further detail below with reference to fig. 1 to 3.
An aeroengine turbine guide vane platform cooling structure comprising:
a flange plate 1 having a cavity therein, the outer side wall of which has a plurality of rows of impingement holes 3 and the inner side wall of which has a plurality of rows of film holes 4;
the multi-row turbulence column 2 is arranged in the cavity and supported between the outer side wall and the inner side wall;
wherein the content of the first and second substances,
each row of the impact holes 3 are communicated with the cavity and distributed in a staggered way;
each exhaust membrane hole 4 is communicated with the cavity and distributed in a staggered way;
the drain flow columns 2 are distributed in a staggered manner;
each impact hole 3 is surrounded by four turbulence columns 2;
each air film hole 4 is surrounded by four turbulence columns 2;
a turbulent flow column 2 is arranged between the adjacent impact holes 3 and the air film holes 4 at intervals.
For the aeroengine turbine guide blade edge plate cooling structure disclosed in the above embodiment, it can be understood by those skilled in the art that, when the aeroengine is in operation, cooling gas may be introduced into the cavity through each of the impingement holes 3, the cooling gas entering the cavity may perform impingement cooling on the inner side wall of the edge plate 1 to form turbulent flow, a counter-vortex system is formed in the middle of the corresponding turbulence columns 2 to form channel vortices with opposite rotation directions, so that sufficient turbulent heat exchange may be performed in the cavity, and the outer side wall of the edge plate 1 may be impacted by swirling to perform convective disturbance heat exchange with the outer side wall, thereby increasing the effective heat exchange area of the cooling gas in the cavity, and finally the cooling gas may be discharged through each of the exhaust film holes to form a gas film on the inner side surface of the edge plate 1, thereby protecting the edge plate 1 from high temperature damage, having a higher cooling efficiency for the edge plate 1, and greatly reducing the consumption of the cooling gas.
For the aero-engine turbine guide vane rim plate cooling structure disclosed in the above embodiment, it can be further understood by those skilled in the art that the rows of impingement holes 3, the film holes 4, and the film holes 2 are arranged in a staggered manner, and each impingement hole 3 and each film hole 4 thereof are surrounded by four film holes 2, and there is a turbulence column 2 between adjacent impingement holes 3 and film holes 4, so as to ensure that the cooling air entering the cavity can be sufficiently turbulent, and the cooling air discharged from each film hole 4 can have a large effective coverage area on the inner side surface of the rim plate 1, so that the cooling efficiency of the rim plate 1 can be further improved, and the consumption of the cooling air can be reduced.
In some optional embodiments, in the aeroengine turbine guide vane platform cooling structure, the diameter of the impingement holes 3 is 1.75 to 2.75 times that of the film holes 4;
the diameter of turbulent flow post 2 is 2.5 times in 4 apertures in the film hole, and the aperture in 4 in the film hole can be confirmed according to concrete reality by relevant technical personnel when using the technical scheme that this application discloses to can make the abundant torrent of the cooling gas that gets into the cavity, increase the circulation ability, reduce the pressure loss to the cooling gas exhaust, and make and cover 1 interior side surface in the listrium on a large scale from the cooling gas of each exhaust film hole 4 and aim at.
In some optional embodiments, in the aeroengine turbine guide blade rim plate cooling structure, the distance between the turbulence columns 2 is 6 to 7.5 times of the aperture of the film holes 4, and the distance between the impact holes 3 and the distance between the film holes 4 are twice of the distance between the turbulence columns 2, so as to ensure that the cooling air entering the cavity can be sufficiently turbulent.
In some alternative embodiments, in the above-mentioned aeroengine turbine guide vane platform cooling structure, the film holes 4 in the diagonal region of the platform 1 have a high density distribution, and the corresponding impingement holes 3 and turbulators 2 also have a high density distribution.
The aeroengine turbine guide blade edge plate is integrally in a rhombic shape, particularly the edge plate of a large chord length turbine guide blade is in a slender rhombic shape, when the aeroengine works, the short diagonal area 5 of the aeroengine turbine guide blade can form the highest local temperature, and meanwhile, the long diagonal area 6 of the aeroengine turbine guide blade edge plate is difficult to cool.
In some alternative embodiments, in the cooling structure of the aeroengine turbine guide vane platform, the distance between the cavity and the edge of the platform 1 is 3 times the diameter of the film hole 4.
In some optional embodiments, in the above-mentioned aeroengine turbine guide vane platform cooling structure, each exhaust film hole 4 is inclined towards the rear end of the platform 1, so that the cooling air discharged from each exhaust film hole 4 flows along the direction of the air flow in the turbine, an air film can be easily formed on the inner side surface of the platform 1 along the direction of the air flow in the turbine, the platform 1 is protected from high temperature damage, interference with the air flow in the turbine can be reduced, and the cooling air can smoothly merge into the air flow in the turbine at the rear end of the platform 1.
In some alternative embodiments, in the above-mentioned aircraft engine turbine guide vane platform cooling structure, the outlet of each exhaust film hole 4 is expanded towards the rear end of the platform 1, and may be in a fan shape, a trapezoid shape, or the like, so as to increase the flow capacity, reduce the pressure loss of the exhaust cooling air, and increase the effective coverage area of the cooling air exhausted from each exhaust film hole 4 on the inner side surface of the platform 1.
In some alternative embodiments, in the above-mentioned aeroengine turbine guide vane platform cooling structure, the diameter of each turbulence column 2 gradually shrinks from the middle portion to the two ends, that is, the portion connected with the inner and outer side walls of the platform 1 has a smaller radial dimension, so that the resistance to the flow of the cooling air in the cavity can be reduced, and the pressure loss can be reduced.
The embodiments are described in a progressive mode in the specification, the emphasis of each embodiment is on the difference from the other embodiments, and the same and similar parts among the embodiments can be referred to each other.
Having thus described the present application in connection with the preferred embodiments illustrated in the accompanying drawings, it will be understood by those skilled in the art that the scope of the present application is not limited to those specific embodiments, and that equivalent modifications or substitutions of related technical features may be made by those skilled in the art without departing from the principle of the present application, and those modifications or substitutions will fall within the scope of the present application.
Claims (8)
1. An aeroengine turbine guide vane platform cooling structure, comprising:
a flange plate (1) having a cavity therein, the outer side wall of which has a plurality of rows of impingement holes (3), and the inner side wall of which has a plurality of rows of film holes (4);
the multi-row turbulence column (2) is arranged in the cavity and supported between the outer side wall and the inner side wall;
wherein, the first and the second end of the pipe are connected with each other,
each row of the impact holes (3) are communicated with the cavity and are distributed in a staggered manner;
each row of the air film holes (4) are communicated with the cavity and distributed in a staggered manner;
each row of the turbulence columns (2) are distributed in a staggered way;
each impact hole (3) is surrounded by four turbulence columns (2);
each air film hole (4) is surrounded by four turbulence columns (2);
and one turbulence column (2) is arranged between the impact hole (3) and the air film hole (4) which are adjacent to each other.
2. The aero engine turbine guide vane platform cooling structure of claim 1,
the aperture of the impact hole (3) is 1.75-2.75 times of the aperture of the air film hole (4);
the diameter of the turbulence column (2) is 2.5 times of the diameter of the air film hole (4).
3. The aero engine turbine guide vane platform cooling structure of claim 1,
the distance between the flow disturbing columns (2) is 6-7.5 times of the diameter of the air film hole (4).
4. The aero engine turbine guide blade platform cooling structure of claim 1,
the diagonal area of the edge plate (1) has a high density distribution of the film holes (4).
5. The aero engine turbine guide blade platform cooling structure of claim 1,
the distance between the cavity and the edge of the edge plate (1) is 3 times of the aperture of the air film hole (4).
6. The aero engine turbine guide vane platform cooling structure of claim 1,
each row of the air film holes (4) is inclined towards the rear end of the edge plate (1).
7. The aero engine turbine guide vane platform cooling structure of claim 1,
the outlet of each row of the air film holes (4) expands towards the rear end of the edge plate (1).
8. The aero engine turbine guide blade platform cooling structure of claim 1,
the diameters of the turbulence columns (2) gradually shrink from the middle part to the two ends.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202211421825.1A CN115492643B (en) | 2022-11-15 | 2022-11-15 | Aeroengine turbine guide blade flange plate cooling structure |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202211421825.1A CN115492643B (en) | 2022-11-15 | 2022-11-15 | Aeroengine turbine guide blade flange plate cooling structure |
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| Publication Number | Publication Date |
|---|---|
| CN115492643A true CN115492643A (en) | 2022-12-20 |
| CN115492643B CN115492643B (en) | 2023-03-14 |
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| CN202211421825.1A Active CN115492643B (en) | 2022-11-15 | 2022-11-15 | Aeroengine turbine guide blade flange plate cooling structure |
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Citations (9)
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| US7921654B1 (en) * | 2007-09-07 | 2011-04-12 | Florida Turbine Technologies, Inc. | Cooled turbine stator vane |
| CN203796330U (en) * | 2014-04-03 | 2014-08-27 | 中国科学院工程热物理研究所 | Cross-arrangement type double-laminate cooling structure |
| CN204024723U (en) * | 2014-08-17 | 2014-12-17 | 中国航空工业集团公司沈阳发动机设计研究所 | A kind of split type laminate cooling structure of turborotor |
| CN113309578A (en) * | 2021-03-22 | 2021-08-27 | 南京航空航天大学 | Novel trough of belt turbulent flow post structure |
| CN113374536A (en) * | 2021-06-09 | 2021-09-10 | 中国航发湖南动力机械研究所 | Gas turbine guide vane |
| CN114109516A (en) * | 2021-11-12 | 2022-03-01 | 中国航发沈阳发动机研究所 | Turbine blade end wall cooling structure |
| US20220127964A1 (en) * | 2020-10-23 | 2022-04-28 | Doosan Heavy Industries & Construction Co., Ltd. | Cooling structure for trailing edge of turbine blade |
| CN115013074A (en) * | 2022-07-26 | 2022-09-06 | 中国航发沈阳发动机研究所 | Cooling structure for upper edge plate of turbine blade of aero-engine |
| CN115163204A (en) * | 2022-08-15 | 2022-10-11 | 中国航发沈阳发动机研究所 | Aeroengine high pressure turbine cascade end wall cooling structure |
-
2022
- 2022-11-15 CN CN202211421825.1A patent/CN115492643B/en active Active
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| US7921654B1 (en) * | 2007-09-07 | 2011-04-12 | Florida Turbine Technologies, Inc. | Cooled turbine stator vane |
| CN203796330U (en) * | 2014-04-03 | 2014-08-27 | 中国科学院工程热物理研究所 | Cross-arrangement type double-laminate cooling structure |
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| CN113309578A (en) * | 2021-03-22 | 2021-08-27 | 南京航空航天大学 | Novel trough of belt turbulent flow post structure |
| CN113374536A (en) * | 2021-06-09 | 2021-09-10 | 中国航发湖南动力机械研究所 | Gas turbine guide vane |
| CN114109516A (en) * | 2021-11-12 | 2022-03-01 | 中国航发沈阳发动机研究所 | Turbine blade end wall cooling structure |
| CN115013074A (en) * | 2022-07-26 | 2022-09-06 | 中国航发沈阳发动机研究所 | Cooling structure for upper edge plate of turbine blade of aero-engine |
| CN115163204A (en) * | 2022-08-15 | 2022-10-11 | 中国航发沈阳发动机研究所 | Aeroengine high pressure turbine cascade end wall cooling structure |
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