CN110925028B - Gas turbine blade with S-shaped impingement cavity partition - Google Patents
Gas turbine blade with S-shaped impingement cavity partition Download PDFInfo
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- CN110925028B CN110925028B CN201911236263.1A CN201911236263A CN110925028B CN 110925028 B CN110925028 B CN 110925028B CN 201911236263 A CN201911236263 A CN 201911236263A CN 110925028 B CN110925028 B CN 110925028B
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- impingement
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- wall surface
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- gas turbine
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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
- F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
- F01D5/12—Blades
- F01D5/14—Form or construction
- F01D5/18—Hollow blades, i.e. blades with cooling or heating channels or cavities; Heating, heat-insulating or cooling means on blades
- F01D5/187—Convection cooling
- F01D5/188—Convection cooling with an insert in the blade cavity to guide the cooling fluid, e.g. forming a separation wall
- F01D5/189—Convection cooling with an insert in the blade cavity to guide the cooling fluid, e.g. forming a separation wall the insert having a tubular cross-section, e.g. airfoil shape
Abstract
The invention relates to a gas turbine blade with an S-shaped impact cavity partition plate, which comprises a blade base body (1), an impact cooling separation cavity (2) positioned in the blade base body and a cold air supply cavity (3) in the middle of the blade, wherein the front side wall surface (7) of the separation cavity of the impact cooling separation cavity (2) is S-shaped; the S-shaped wall surface restrains the cold air impacting the target area, so that the cold air performs sequential diffusion flow along an expansion channel formed by restraining the S-shaped wall surface, and the cold air is distributed more uniformly; the S-shaped wall surface is restrained, so that cold air is uniformly filled, the cooling effect is uniform, and the temperature difference of the wall surface is overlarge; the S-shaped wall surface is combined with the turbulence ribs to ensure that the cold air is distributed more uniformly, and the heat exchange effect is further improved.
Description
Technical Field
The invention relates to a cooling structure for high-temperature components of an aircraft engine and a gas turbine, which is particularly suitable for high-temperature turbine blades.
Background
The turbine front temperature of the high-performance aircraft engine and the gas turbine reaches more than 2000K at present, and far exceeds the long-term allowable temperature of metal materials of turbine blades. In order to ensure the safety of the turbine blade, a complicated cooling structure is used inside the blade, and the blade is cooled by cooling air.
In a 2000K inlet temperature class turbine blade design, cooling air is typically forced through impingement holes in the interior of the blade to impinge on the cell wall surfaces in the interior of the turbine blade. The heat exchange in the blade is strengthened through the impact effect of cold air, so that more heat conducted from the outside of the blade to the inside is taken away, and the metal matrix of the blade is ensured to be at a lower working temperature. But because the cold air strikes regional limited, in the target area of impact, the heat transfer effect reinforcing, but keeps away from the position of target area, and the heat transfer effect is relatively poor, and the cold air cover surface is little, leads to the inhomogeneity enhancement of the metal matrix temperature field of blade, and then makes turbine blade reliability and life-span under the high temperature condition reduce. Meanwhile, due to the fact that the flow in the impact separation chamber is disordered, the flow loss is large, and the subsequent flow of cold air is further influenced. For avoiding the occurrence of this problem, the flow of the cold air after impacting the wall surface needs to be restrained and arranged reasonably, so that the coverage of the cold air is wider, the distribution of the cold air is more uniform, and the flow loss of the cold air is smaller. Therefore, the invention provides a novel cooling structure, which is a long and narrow partition chamber positioned in the blade, and the impact cold air is guided through the S-shaped wall surface on one side of the partition chamber, so that the cold air is uniformly distributed in the partition chamber, and the problem of the traditional design is solved.
Disclosure of Invention
The purpose of the invention is as follows:
the invention aims to provide a cooling structure with higher cooling effect and capability of reducing temperature difference, aiming at the problems of uneven cold air distribution, large cooling effect difference and large wall temperature difference of the currently adopted impingement cooling partition cavity structure.
The invention realizes the scheme of the above purpose:
the invention provides a gas turbine blade with an S-shaped impingement cavity partition plate, which comprises a blade base body 1, an impingement cooling partition cavity 2 positioned in the blade base body, and a cold air supply cavity 3 in the middle of the blade, wherein the front side wall 7 of the partition cavity of the impingement cooling partition cavity 2 is S-shaped.
Preferably, in the gas turbine blade with the S-shaped impingement cavity partition, the impingement cooling compartment 2 has impingement holes 5 on the inner side wall surface 4 of the compartment, and the impingement holes 5 are connected with the cold air supply cavity 3.
Preferably, in the gas turbine blade with the S-shaped impingement cavity partition, the aft side of the impingement cooling compartment 2 is the compartment outlet.
Preferably, in a gas turbine blade with an S-shaped impingement cavity partition, the aft side of the impingement cooling compartment 2 is a solid wall surface, which may comprise the compartment forward side wall surface 7 or another wall surface of the next impingement cooling compartment 2.
Preferably, in the gas turbine blade with the S-shaped impingement cavity partition, the outer wall surface 6 of the impingement cooling compartment 2 is provided with film holes 9, and the film holes 9 are located downstream of the position 10 on the outer wall surface 6 of the compartment opposite to the impingement holes 5.
Preferably, in the gas turbine blade with the S-shaped impact cavity partition plate, the turbulence ribs 11 are arranged between the film holes 9 and positions 10 on the outer wall surface 6 of the compartment opposite to the impact holes 5.
Preferably, the gas turbine blade with the S-shaped impact cavity partition plate is provided with 2-5 rows of the turbulence ribs 11, and the turbulence ribs 11 are approximately arc-shaped.
Preferably, the turbulence ribs 11 corresponding to the plurality of impingement holes 5 of the gas turbine blade with the S-shaped impingement cavity partition plate are connected at the last row to form a complete curve consisting of a plurality of approximate circular arcs along the whole blade height direction.
Preferably, in the gas turbine blade with the S-shaped impact cavity partition plate, the plurality of impact holes 5 are arranged on the inner side wall surface 4 of the separation cavity in the middle of the blade along the blade height direction, and both ends of the blade are located at the downstream of the impact holes 5.
Preferably, the gas turbine blade with the S-shaped impingement cavity partition is provided with a plurality of bypass flow columns 12 in the area between the film holes 9 and the turbulence ribs 11.
Preferably, the Mach number of the outlet of the impingement hole 5 of the gas turbine blade with the S-shaped impingement cavity partition is 0.3-0.5.
Preferably, in the gas turbine blade with the S-shaped impact cavity partition plate, the distance between the outer wall surface 6 of the compartment and the inner wall surface 4 of the compartment is 1-3 times of the diameter of the impact hole 5.
Preferably, the height of the turbulence rib 11 of the gas turbine blade with the S-shaped impact cavity partition plate is about 0.2-0.5 times of the diameter of the impact hole 5.
The invention has the beneficial effects that:
the gas turbine blade with the S-shaped impact cavity partition plate has the advantages that: (1) the S-shaped wall surface restrains the cold air impacting the target area, so that the cold air performs sequential diffusion flow along an expansion channel formed by restraining the S-shaped wall surface, and the cold air is distributed more uniformly; (2) the S-shaped wall surface is restrained, so that cold air is uniformly filled, the cooling effect is uniform, and the temperature difference of the wall surface is overlarge; (3) the S-shaped wall surface is combined with the turbulence ribs to ensure that the cold air is distributed more uniformly, and the heat exchange effect is further improved.
Drawings
FIG. 1 illustrates a currently common internal impingement cavity of a turbine bucket;
FIG. 2 is a turbine blade with an S-shaped impingement cavity diaphragm;
FIG. 3 is an impingement cavity configuration of a turbine bucket with an S-shaped impingement cavity diaphragm;
FIG. 4 illustrates an arrangement of turbulator ribs on the outer wall of a turbine blade with an S-shaped impingement cavity baffle;
FIG. 5 is an expanded application of a turbine blade with an S-shaped impingement cavity diaphragm;
FIG. 6 is a schematic view of the effect;
1-a blade base body; 2-impingement cooling of the cell; 3-a cold air supply cavity; 4-inner side wall surface of the partition cavity; 5-impact holes; 6-outer wall surface of the separate cavity; a 7-S shaped compartment front sidewall face; 8-the rear side wall surface of the compartment or the compartment outlet; 9-air film hole; 10-impact holes are arranged at corresponding positions on the outer side wall surface of the separate cavity; 11-a spoiler rib; 12-a bypass column; 13-dead space where cold air cannot reach.
Detailed Description
In order to make the objects, technical solutions and advantages of the present invention more apparent, the present invention is described in further detail below with reference to specific embodiments and the accompanying drawings.
The invention is described in detail below with reference to the accompanying drawings and specific embodiments.
A gas turbine blade with an S-shaped impingement cavity barrier includes a blade base 1, an impingement cooling compartment 2 located within the blade base. The inner wall surface 4 of the compartment has impingement holes 5 which are connected to the cold air supply chamber 3 inside the blade. The compartment outer wall 6 is the outer surface of the vane. The front side wall 7 of the compartment close to the impact hole is S-shaped, and the rear side 8 is the compartment outlet or solid wall. When the rear side is a solid wall the cold air flows out of the compartment through the film holes 9 in the outer side wall 6 of the compartment.
2-5 rows of turbulence ribs 11 are arranged at the downstream of a position 10 on the outer side wall surface 6 of the separation cavity opposite to each impact hole 5, the turbulence ribs 11 are approximately arc-shaped, and the turbulence ribs 11 corresponding to each impact hole 5 are connected into a complete turbulence rib 11 along the whole blade height direction at the position of the last row of turbulence ribs 11.
The diameter of the impingement hole 5 of the gas turbine blade with the S-shaped impingement cavity partition is about 1-3mm, and the important point is to ensure that the Mach number of the exit of the impingement hole is about 0.3-0.5. The distance between the outer wall surface 6 of the compartment and the inner wall surface 4 of the compartment is about 1-3 times the diameter of the impingement holes. The interval of the turbulence ribs 11 is about 2-4 times of the impact aperture, and the height of the turbulence ribs 11 is about 0.2-0.5 times of the impact aperture.
The cooling structure can be combined in multiple groups, and then is used for the internal cooling design of the whole blade, and can be combined with the turbulence column 12 to be used according to the use conditions in the impact cavity, so that a better cooling effect is obtained.
In the existing design, the linear impact cavity partition plate shown in fig. 1 is adopted, and the uniformly distributed impact holes form corresponding circular target areas on the outer wall surface 6 of the separate cavity. Between the circular target areas, a dead zone 13 is formed, which is not reached by the cold air, and this results in a low flow rate of the cold air, a small coverage of the cold air, and a poor cooling effect in this region. The invention is based on the feature of the flow that the borderlines of the dead space are replaced by S-shaped walls, thus allowing cold air to fill the whole compartment after impacting the target surface. And meanwhile, under the guidance of the S-shaped wall surface, the flow is gradually diffused, so that the flow loss in the cavity is reduced. Thereby obtaining uniform cold air covering and cooling effect and generating uniform temperature field.
The above description is only for the purpose of illustrating the preferred embodiments of the present invention and is not to be construed as limiting the invention, and any modifications, equivalents and improvements made within the spirit and principle of the present invention are intended to be included within the scope of the present invention. The above description is only a preferred embodiment of the present invention, and it should be noted that, for those skilled in the art, several modifications and variations can be made without departing from the technical principle of the present invention, and these modifications and variations should also be regarded as the protection scope of the present invention.
Claims (9)
1. A gas turbine blade with an S-shaped impingement cavity partition plate comprises a blade base body (1), an impingement cooling separation cavity (2) located inside the blade base body and a cold air supply cavity (3) in the middle of the blade, and is characterized in that the front side wall surface (7) of the separation cavity of the impingement cooling separation cavity (2) is S-shaped;
the inner side wall surface (4) of the separate cavity of the impingement cooling separate cavity (2) is provided with impingement holes (5), and the impingement holes (5) are connected with the cold air supply cavity (3);
the outer wall surface (6) of the separate cavity of the impingement cooling separate cavity (2) is provided with a gas film hole (9), and the gas film hole (9) is positioned at the downstream of a position (10) on the outer wall surface (6) of the separate cavity opposite to the impingement hole (5);
turbulence ribs (11) are arranged between the air film holes (9) and positions (10) on the outer wall surface (6) of the separation cavity opposite to the impact holes (5);
the turbulence ribs (11) are approximately arc-shaped, and the turbulence ribs (11) corresponding to the impact holes (5) are connected into a complete curve formed by a plurality of approximately arc-shaped ribs along the whole blade height direction at the last row.
2. Gas turbine blade with S-shaped impingement cavity barrier according to claim 1, characterised in that the aft side of the impingement cooling compartment (2) is the compartment outlet.
3. Gas turbine blade with S-shaped impingement cavity partition according to claim 1, characterised in that the aft side of the impingement cooling compartment (2) is a solid wall surface, which solid wall surface comprises the compartment forward side wall surface (7) or other wall surface of the next impingement cooling compartment (2).
4. The gas turbine blade with S-shaped impingement cavity divider as claimed in claim 1, wherein said turbulator ribs (11) have 2-5 rows.
5. The gas turbine blade with an S-shaped impingement cavity partition according to claim 4, characterized in that several impingement holes (5) are arranged in the vane height direction on the inner sidewall surface (4) of the partition in the middle of the blade, both ends of the blade being downstream of the impingement holes (5).
6. Gas turbine blade with S-shaped impingement cavity diaphragm according to claim 5, characterised in that several turbulence columns (12) are arranged in the area between the film hole (9) and the turbulence ribs (11).
7. The gas turbine blade with S-shaped impingement cavity partitions of claim 6, characterized in that the exit of the impingement holes (5) has a mach number of 0.3-0.5.
8. A gas turbine blade with S-shaped impingement cavity partition according to claim 1, characterised in that the distance between the outer compartment wall surface (6) and the inner compartment wall surface (4) is 1-3 times the diameter of the impingement holes (5).
9. The gas turbine blade with S-shaped impingement cavity partition according to claim 1, characterized in that the height of the turbulator ribs (11) is 0.2-0.5 times the diameter of the impingement hole (5).
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN201911236263.1A CN110925028B (en) | 2019-12-05 | 2019-12-05 | Gas turbine blade with S-shaped impingement cavity partition |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN201911236263.1A CN110925028B (en) | 2019-12-05 | 2019-12-05 | Gas turbine blade with S-shaped impingement cavity partition |
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| Publication Number | Publication Date |
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| CN110925028A CN110925028A (en) | 2020-03-27 |
| CN110925028B true CN110925028B (en) | 2022-06-07 |
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| CN201911236263.1A Active CN110925028B (en) | 2019-12-05 | 2019-12-05 | Gas turbine blade with S-shaped impingement cavity partition |
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Families Citing this family (6)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN112145234B (en) * | 2020-09-24 | 2021-08-20 | 大连理工大学 | Omega type gyration chamber plywood cooling structure |
| 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 |
| CN114198154B (en) * | 2021-12-15 | 2023-08-15 | 中国科学院工程热物理研究所 | Cooling structure |
| CN114526125B (en) * | 2022-04-24 | 2022-07-26 | 中国航发四川燃气涡轮研究院 | Cooling unit with rotary cavity for bag and turbine blade structure |
| CN115013076B (en) * | 2022-08-10 | 2022-10-25 | 中国航发四川燃气涡轮研究院 | Gondola water faucet form turbine blade cooling unit and turbine blade |
| CN115045721B (en) * | 2022-08-17 | 2022-12-06 | 中国航发四川燃气涡轮研究院 | Series-type rotational flow impact turbine blade cooling unit and turbine blade |
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| CN1477292A (en) * | 2002-07-11 | 2004-02-25 | �����ع�ҵ��ʽ���� | Turbomachine blade and gas turbomachine |
| CN1786426A (en) * | 2005-12-26 | 2006-06-14 | 北京航空航天大学 | Pulsing impact cooling blade for gas turbine engine |
| US8777569B1 (en) * | 2011-03-16 | 2014-07-15 | Florida Turbine Technologies, Inc. | Turbine vane with impingement cooling insert |
| CN205382958U (en) * | 2016-03-02 | 2016-07-13 | 中航商用航空发动机有限责任公司 | Turbine blade and aeroengine |
| CN106437862A (en) * | 2015-07-29 | 2017-02-22 | 安萨尔多能源英国知识产权有限公司 | Method for cooling a turbo-engine component and turbo-engine component |
Family Cites Families (1)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| EP3232002A1 (en) * | 2016-04-11 | 2017-10-18 | Rolls-Royce Corporation | Impingement plate with stress relief feature |
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2019
- 2019-12-05 CN CN201911236263.1A patent/CN110925028B/en active Active
Patent Citations (5)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN1477292A (en) * | 2002-07-11 | 2004-02-25 | �����ع�ҵ��ʽ���� | Turbomachine blade and gas turbomachine |
| CN1786426A (en) * | 2005-12-26 | 2006-06-14 | 北京航空航天大学 | Pulsing impact cooling blade for gas turbine engine |
| US8777569B1 (en) * | 2011-03-16 | 2014-07-15 | Florida Turbine Technologies, Inc. | Turbine vane with impingement cooling insert |
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| CN110925028A (en) | 2020-03-27 |
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