CN113090335A - Impact air-entraining film double-wall cooling structure for turbine rotor blade - Google Patents
Impact air-entraining film double-wall cooling structure for turbine rotor blade Download PDFInfo
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- CN113090335A CN113090335A CN202110526654.8A CN202110526654A CN113090335A CN 113090335 A CN113090335 A CN 113090335A CN 202110526654 A CN202110526654 A CN 202110526654A CN 113090335 A CN113090335 A CN 113090335A
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- 238000001816 cooling Methods 0.000 title claims abstract description 51
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- 238000010586 diagram Methods 0.000 description 6
- 239000002131 composite material Substances 0.000 description 3
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- 238000003491 array Methods 0.000 description 2
- 230000005764 inhibitory process Effects 0.000 description 2
- 230000009286 beneficial effect Effects 0.000 description 1
- 230000006866 deterioration Effects 0.000 description 1
- 239000002737 fuel gas Substances 0.000 description 1
- 230000002401 inhibitory effect Effects 0.000 description 1
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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
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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
- F01D5/186—Film cooling
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- Turbine Rotor Nozzle Sealing (AREA)
Abstract
The invention discloses an impact air-entrainment film double-wall cooling structure for a turbine rotor blade, wherein the rotor blade comprises a front edge impact area and a tail edge impact area, the front edge impact area comprises a front edge air supply inner cavity and a front edge impact inner cavity which are positioned in the middle of the rotor blade, and the front edge impact area also comprises a first blade back double-wall cavity, a second blade back double-wall cavity and a first blade basin double-wall cavity, the first blade back double-wall cavity is arranged in a blade back area of the rotor blade, and the first blade basin double-wall cavity is arranged in a blade basin area of the rotor blade; the trailing edge impact area comprises a trailing edge air supply inner cavity, a transverse turbulence column impact cavity and a trailing edge exhaust cavity which are positioned in the middle of the rotor blade, and the trailing edge impact area also comprises a third blade back double-layer wall cavity arranged in the blade back area of the rotor blade and a second blade basin double-layer wall cavity arranged in the blade basin area of the rotor blade.
Description
Technical Field
The invention belongs to the technical field of turbine blades of aero-engines and gas turbines, and particularly relates to an impact air entrainment film double-wall cooling structure for a turbine rotor blade.
Background
With the development of new-generation aircraft engines and gas turbines, the thermal cycle parameters of the engines are continuously improved, and the thermal protection requirement of turbine blades as key hot-end components is increasingly outstanding. The thermal protection of the turbine blade mainly adopts an air cooling technology, and the normal work of the blade in a high-temperature environment is met by designing a complicated cold air flow form inside. The basic cooling forms of the prior aeroengine turbine blade comprise convection cooling, impingement cooling, film cooling and the like, and most of the cooling blades adopt a composite cooling design of one or more cooling forms. Among them, the combination of convection cooling and film cooling, and the combination of impingement cooling and film cooling are the most commonly used composite cooling methods.
The cooling of the leading edge area of the turbine rotor blade under the existing high thermodynamic cycle parameters mainly adopts an impact air-entrainment film cooling mode, and the cold air impact cooling of the leading edge area is realized by air supply of a radial flowing ribbed channel. The cooling of the middle part of the blade body mainly adopts a multi-cavity ribbed convection cooling channel air-entraining film composite cooling mode. The technical scheme is mature and applied, and the heat exchange capacity of the inner cavity can be improved through the design of the cooling structure of the rib-type channel turbulence ribs. However, with the increasing thermal load of the blades, it has become increasingly difficult for such a combined cooling method of ribbed channel air-film cooling to meet the cooling requirements of the turbine rotor blades.
The main disadvantages of the above technical solutions are represented in the following three aspects: on one hand, the heat exchange capacity of the ribbed channels in the front edge air supply cavity and the middle part of the blade body is improved to a small extent, and a large amount of cold air and larger pressure loss of the cold air are needed to improve the heat exchange coefficient of the inner cavity; secondly, the ribbed channel cannot effectively cool a high-heat-load area after the fuel gas side transition of the blade back, and a local high-temperature area of the blade is easy to appear; thirdly, the large-cavity ribbed channels in the rotor blades are obviously influenced by the rotation effect, and the heat exchange capacity of the suction side is obviously reduced under the negative inhibition effect of the rotation, so that the heat exchange capacity of the inner cavity is further deteriorated.
Disclosure of Invention
The invention aims to provide an impact air-entrainment film double-wall cooling structure for a turbine rotor blade, which has higher inner cavity heat exchange capacity and very high cooling effect compared with a conventional ribbed channel combined cooling structure, can realize the optimized design based on heat management and the negative effect of rotation inhibition, and can meet the requirement of thermal protection of the turbine rotor blade at higher turbine front inlet temperature.
The purpose of the invention can be realized by the following technical scheme:
an impact air entrainment film double-wall cooling structure for a turbine rotor blade, the rotor blade comprises a front edge impact area and a tail edge impact area, the front edge impact area comprises a front edge air supply inner cavity and a front edge impact inner cavity which are positioned in the middle of the rotor blade, the front edge impact area also comprises a first blade back double-wall cavity, a second blade back double-wall cavity and a first blade basin double-wall cavity, the first blade back double-wall cavity and the second blade back double-wall cavity are arranged in a blade back area of the rotor blade, and the first blade basin double-wall cavity is arranged in a blade basin area; the trailing edge strikes district and strikes chamber and trailing edge exhaust chamber including trailing edge air feed inner chamber, horizontal vortex post that are located rotor blade middle part, trailing edge strikes district still including setting up at the regional third leaf back double-layer wall chamber of rotor blade leaf back of the body and setting up at the regional second leaf basin double-layer wall chamber of rotor blade leaf basin.
As a further scheme of the invention: and the side wall of the front edge air supply inner cavity is respectively provided with a front edge air film hole, a blade back cheek area air film hole and a blade basin cheek area air film hole.
As a further scheme of the invention: and front edge inner cavity impact holes communicated with each other are arranged between the front edge air supply inner cavity and the front edge impact inner cavity.
As a further scheme of the invention: the cavity walls of the first blade back double-layer wall cavity and the second blade back double-layer wall cavity are both provided with blade back double-layer wall air film holes, and the cavity wall of the first blade basin double-layer wall cavity is provided with a blade basin double-layer wall air film hole.
As a further scheme of the invention: the first blade back double-layer wall cavity and the second blade back double-layer wall cavity are respectively provided with a first blade back impact hole and a second blade back impact hole, and the first blade back impact hole and the second blade back impact hole are communicated with the front edge impact inner cavity; the first blade basin double-layer wall cavity is provided with a first blade basin impact hole, and the first blade basin impact hole is communicated with the front edge impact inner cavity.
As a further scheme of the invention: and a trailing edge inner cavity impact hole is arranged between the trailing edge air supply inner cavity and the transverse turbulence column impact cavity and communicated with each other.
As a further scheme of the invention: and a third blade back impact hole is formed in the third blade back double-layer wall cavity and communicated with the tail edge air supply inner cavity.
As a further scheme of the invention: and a leaf basin trailing edge air film hole and a second leaf basin impact hole are arranged at the position of the double-layer wall cavity of the second leaf basin, the leaf basin trailing edge air film hole is communicated with the outside of the rotor blade, and the second leaf basin impact hole is communicated with a trailing edge air supply inner cavity.
As a further scheme of the invention: the transverse flow disturbing column impact cavity, the third vane back double-layer wall cavity and the second vane basin double-layer wall cavity are provided with outflow seams communicated with a tail edge exhaust cavity, and the tail edge exhaust cavity is provided with a tail edge middle cleft seam.
As a further scheme of the invention: and a plurality of groups of blade back double-layer wall cavity flow disturbing columns are arranged in the third blade back double-layer wall cavity.
As a further scheme of the invention: and a plurality of groups of flow disturbing columns with double-layer wall cavities of the leaf basins are arranged in the double-layer wall cavities of the second leaf basin.
As a further scheme of the invention: and a plurality of groups of transverse inner cavity turbulence columns are arranged in the transverse turbulence column impact cavity.
As a further scheme of the invention: and a plurality of groups of turbulence columns of the exhaust inner cavity are arranged in the tail edge exhaust cavity.
The invention has the beneficial effects that:
1) the double-wall type impingement cooling of the inner cavities of the middle part and the tail edge of the blade can realize higher heat exchange coefficient of the inner cavity, and the cooling effect of the blade is better;
2) the cold air can be throttled through the impact hole, the throttling adjustment range of the cold air is large, and the design flexibility is high;
3) the impingement cooling form has a strong inhibiting effect on the negative effect induced by rotation, and the phenomenon of local heat exchange deterioration on the rotor piece is not easy to occur.
Drawings
In order to more clearly illustrate the technical solutions of the embodiments of the present invention, the drawings used in the description of the embodiments will be briefly introduced below, and it is obvious that the drawings in the following description are only some embodiments of the present invention, and it is obvious for those skilled in the art that other drawings can be obtained according to the drawings without creative efforts.
FIG. 1 is a schematic view of the overall structure of the present invention;
FIG. 2 is a schematic diagram of the distribution structure of the inner cavity of the present invention;
FIG. 3 is a schematic diagram of the distribution structure of the gas film holes of the present invention;
FIG. 4 is a schematic diagram of the distribution of the impingement holes of the present invention;
FIG. 5 is a schematic diagram of the distribution structure of the turbulence column of the present invention;
FIG. 6 is a schematic diagram of the arrangement of the impact holes and the air film holes of the double-wall zone according to the present invention.
In the figure: 1. a leading edge impingement zone; 2. a trailing edge impingement zone; 3. the leading edge impacts the inner cavity; 4. a front edge air supply cavity; 5. a tail edge air supply cavity; 6. a transverse turbulence column impact cavity; 7. a trailing edge exhaust cavity; 8. a first leaf back double-walled cavity; 9. a second leaf back double-wall cavity; 10. a third leaf-back double-walled cavity; 11. a first leaf basin double-wall cavity; 12. a second double-wall cavity of the second basin; 13. a leading edge air film hole; 14. the air film hole of the blade back cheek area; 15. air film holes in the cheek area of the leaf basin; 16. a leaf back double-wall air film hole; 17. leaf basin double-walled air film pores; 18. a leaf basin trailing edge air film hole; 19. splitting in the tail edge; 20. leading edge inner cavity impingement holes; 21. a first blade back impingement hole; 22. a second leaf back impingement hole; 23. a third lobe back impingement hole; 24. a first bucket impingement hole; 25. a second basin impingement hole; 26. a trailing edge inner cavity impingement hole; 27. a flow disturbing column with a double-layer wall cavity on the blade back; 28. the double-layer wall cavity turbulence column of the leaf basin; 29. a transverse inner cavity turbulence column; 30. the exhaust cavity is provided with a turbulence column.
Detailed Description
The technical solutions of the present invention will be described clearly and completely with reference to the following embodiments, and it should be understood that the described embodiments are only a part of the embodiments of the present invention, and not all of the embodiments. All other embodiments, which can be derived by a person skilled in the art from the embodiments given herein without making any creative effort, shall fall within the protection scope of the present invention.
Referring to fig. 1-6, an impact aerated film double-wall cooling structure for a turbine rotor blade includes a leading edge impact region 1 and a trailing edge impact region 2, where the leading edge impact region 1 includes a leading edge air supply inner cavity 4 and a leading edge impact inner cavity 3 located in the middle of the rotor blade, and the leading edge impact region 1 further includes a first blade back double-wall cavity 8, a second blade back double-wall cavity 9 and a first blade basin double-wall cavity 11 arranged in a blade basin region of the rotor blade; the trailing edge impact area 2 comprises a trailing edge air supply inner cavity 5, a transverse spoiler column impact cavity 6 and a trailing edge exhaust cavity 7 which are positioned in the middle of the rotor blade, and the trailing edge impact area 2 further comprises a third blade back double-layer wall cavity 10 arranged in the rotor blade back area and a second blade basin double-layer wall cavity 12 arranged in the rotor blade basin area.
And the side wall of the front edge air supply inner cavity 4 is respectively provided with a front edge air film hole 13, a blade back cheek area air film hole 14 and a blade basin cheek area air film hole 15.
The front edge air supply inner cavity 4 is communicated with the front edge impact inner cavity 3 through a front edge inner cavity impact hole 20.
The wall of the first blade back double-layer wall cavity 8 and the wall of the second blade back double-layer wall cavity 9 are both provided with blade back double-layer wall air film holes 16, and the wall of the first blade basin double-layer wall cavity 11 is provided with blade basin double-layer wall air film holes 17.
The first blade back double-layer wall cavity 8 and the second blade back double-layer wall cavity 9 are respectively provided with a first blade back impact hole 21 and a second blade back impact hole 22, and the first blade back impact hole 21 and the second blade back impact hole 22 are communicated with the front edge impact inner cavity 3; the first blade basin double-layer wall cavity 11 is provided with a first blade basin impact hole 24, and the first blade basin impact hole 24 is communicated with the front edge impact inner cavity 3.
And a tail edge inner cavity impact hole 26 is arranged between the tail edge air supply inner cavity 5 and the transverse turbulence column impact cavity 6 and communicated with each other.
And a third blade back impact hole 23 is formed in the third blade back double-layer wall cavity 10, and the third blade back impact hole 23 is communicated with the tail edge air supply inner cavity 5.
And a leaf basin trailing edge air film hole 18 and a second leaf basin impact hole 25 are arranged at the position of the second leaf basin double-layer wall cavity 12, the leaf basin trailing edge air film hole 18 is communicated with the outside of the rotor blade, and the second leaf basin impact hole 25 is communicated with the trailing edge air supply inner cavity 5.
The transverse turbulence column impact cavity 6, the third blade back double-layer wall cavity 10 and the second blade basin double-layer wall cavity 12 are all provided with outflow seams communicated with the trailing edge exhaust cavity 7, and the trailing edge exhaust cavity 7 is provided with a trailing edge middle cleft seam 19.
A plurality of groups of blade back double-layer wall cavity flow disturbing columns 27 are arranged in the third blade back double-layer wall cavity 10.
A plurality of groups of flow disturbing columns 28 of the double-layer wall cavity of the leaf basin are arranged in the double-layer wall cavity 12 of the second leaf basin.
And a plurality of groups of transverse inner cavity flow disturbing columns 29 are arranged in the transverse flow disturbing column impact cavity 6.
A plurality of groups of air exhaust inner cavity turbulence columns 30 are arranged in the tail edge air exhaust cavity 7.
When the blade is used, cold air in the front edge air supply inner cavity 4 flows in through the bottom of the blade tenon, and respectively enters the front edge impact inner cavity 3, the first blade back impact hole 21 and the second blade back impact hole 22 through the front edge inner cavity impact hole 20, the first blade back double-layer wall cavity 8, the second blade back double-layer wall cavity 9 and the first blade basin impact hole 24, and enters the first blade basin double-layer wall cavity 11, so that impact cooling is formed on the front edge of the blade, the wall surface of the middle blade basin and the wall surface of the blade back.
After entering the front edge impact inner cavity 3, the cold air flows out through the front edge air film holes 13, the blade back cheek area air film holes 14 and the blade basin cheek area air film holes 15, and forms air film cooling protection for the blade body. The leading edge air film holes 13 can be arranged in a range of 2-5 air film discharge holes according to the diameter of the leading edge of the blade, and the blade back cheek area air film holes 14 and the blade basin cheek area air film holes 15 are generally designed in a row. The cold air impacts the first blade back double-layer wall cavity 8 and the second blade back double-layer wall cavity 9 and then flows out through the single-row blade back double-layer wall air film holes 16 respectively, enters the first blade basin double-layer wall cavity 11 and then flows out through the single-row blade basin double-layer wall air film holes 17, and air film protection is formed in a blade back area and a blade basin area in the middle of the blade respectively.
The front edge inner cavity impingement holes 20 are arranged in a radial single-row equidistant uniform hole mode, and the total flow through area of the impingement holes is not smaller than the total outflow area of the front edge air film holes 13 on the front edge impingement inner cavity 3. The first blade back impact hole 21, the second blade back impact hole 22 and the first blade basin impact hole 24 are all arranged in radial double rows, the arrangement schematic diagram of the impact holes and the air film holes refers to fig. 6, the aperture of the impact holes is gradually reduced from the blade root area to the blade tip area in the radial direction, but the total flow area of the impact holes is larger than the total outflow area of the air film holes.
The cold air of the tail edge air supply inner cavity 5 flows in through the bottom of the blade tenon, and respectively enters the transverse turbulence column impact cavity 6 through the tail edge inner cavity impact hole 26, enters the third blade back double-layer wall cavity 10 through the third blade back impact hole 23, enters the second blade basin double-layer wall cavity 12 through the second blade basin impact hole 25, and forms impact cooling on the blade basin wall surface and the blade back wall surface of the blade middle part and the tail edge area.
The transverse turbulence column impact cavity 6 is a turbulence column cavity, and three rows of turbulence column arrays in a radial staggered form are designed in the spanwise direction. After entering the transverse turbulence column impact cavity 6 through transverse impact of the tail edge air supply inner cavity 5, the cold air enters the tail edge exhaust cavity 7 through the turbulence column array from an outflow seam of the transverse turbulence column impact cavity 6. The third blade back double-layer wall cavity 10 and the second blade basin double-layer wall cavity 12 are impact turbulence column cavities, cold air directly impacts the blade back and the side wall surface of the blade basin through the trailing edge air supply inner cavity 5, and flows into the trailing edge exhaust cavity 7 from the outflow seam of the third blade back double-layer wall cavity 10 and the outflow seam of the second blade basin double-layer wall cavity 12 through the turbulence column array of the double-layer wall cavities. The second blade basin double-layer wall cavity 12 can be provided with 2-3 exhaust film holes on the wall surface according to the thermal load condition outside the blade, and is used for thermal protection of the blade basin side in the tail edge area of the blade. After three strands of cold air are converged into the tail edge exhaust cavity 7, the cold air flows out from the tail edge middle cleft 19 through three rows of radial staggered turbulence column arrays.
The double-wall impact holes connected with the front edge air supply inner cavity 4 and the tail edge air supply inner cavity 5 can realize the adjustment of radial arrangement and spanwise arrangement according to the actual thermal load distribution of the blades, can realize the cooling and heat management design of the blades and have great designability.
The array design of the transverse inner cavity turbulence column 29 and the exhaust inner cavity turbulence column 30 is to enhance the connection rigidity of the blade back side and the blade basin side on one hand, and to enhance the inner cavity heat exchange through the area increase and the turbulence effect of the turbulence columns on the other hand.
The array arrangement of the flow disturbing columns 27 of the blade back double-layer wall cavity and the flow disturbing columns 28 of the blade basin double-layer wall cavity not only enhances the connection rigidity and enhances the heat exchange, but also plays a role in the cold air throttling of the third blade back double-layer wall cavity 10 and the second blade basin double-layer wall cavity 12.
The third blade-back double-layer wall cavity 10 has no air film hole to flow out, all cold air flows into the trailing edge exhaust cavity 7 from the outflow seam of the third blade-back double-layer wall cavity 10, and the air film outflow at the blade back side has a large negative influence on the pneumatic performance of the turbine.
The preferred embodiments of the invention disclosed above are intended to be illustrative only. The preferred embodiments are not intended to be exhaustive or to limit the invention to the precise embodiments disclosed. Obviously, many modifications and variations are possible in light of the above teaching. The embodiments were chosen and described in order to best explain the principles of the invention and the practical application, to thereby enable others skilled in the art to best utilize the invention. The invention is limited only by the claims and their full scope and equivalents.
Claims (13)
1. An impact air entrainment film double-wall cooling structure for a turbine rotor blade, the rotor blade comprising a leading edge impact zone (1) and a trailing edge impact zone (2), characterized in that the leading edge impact zone (1) comprises a leading edge air supply inner cavity (4) and a leading edge impact inner cavity (3) which are located in the middle of the rotor blade, the leading edge impact zone (1) further comprises a first blade back double-wall cavity (8) and a second blade back double-wall cavity (9) which are arranged in the blade back area of the rotor blade and a first blade basin double-wall cavity (11) which is arranged in the blade basin area of the rotor blade; the trailing edge impact area (2) comprises a trailing edge air supply inner cavity (5), a transverse turbulence column impact cavity (6) and a trailing edge exhaust cavity (7) which are positioned in the middle of the rotor blade, and the trailing edge impact area (2) further comprises a third blade back double-layer wall cavity (10) arranged in the rotor blade back area and a second blade basin double-layer wall cavity (12) arranged in the rotor blade basin area.
2. The impingement plenum double wall cooling structure for turbine rotor blades according to claim 1, wherein the leading edge supply cavity (4) side walls are provided with leading edge film holes (13), blade back cheek region film holes (14) and blade basin cheek region film holes (15), respectively.
3. The impingement plenum double wall cooling structure for turbine rotor blades according to claim 1, wherein leading edge cavity impingement holes (20) are provided between the leading edge air supply cavity (4) and the leading edge impingement cavity (3) in communication.
4. The impingement plenum cooling structure for turbine rotor blades according to claim 1, wherein the cavity walls of the first blade back double-walled cavity (8) and the second blade back double-walled cavity (9) are provided with blade back double-walled plenum holes (16), and the cavity wall of the first blade basin double-walled cavity (11) is provided with blade basin double-walled plenum holes (17).
5. The impact air-entrainment film double-wall cooling structure for the turbine rotor blade according to claim 1, characterized in that the first blade back double-wall cavity (8) and the second blade back double-wall cavity (9) are respectively provided with a first blade back impact hole (21) and a second blade back impact hole (22), and the first blade back impact hole (21) and the second blade back impact hole (22) are both communicated with the leading edge impact inner cavity (3); the first blade basin double-layer wall cavity (11) is provided with a first blade basin impact hole (24), and the first blade basin impact hole (24) is communicated with the front edge impact inner cavity (3).
6. The impingement plenum double wall cooling structure for turbine rotor blades according to claim 1, wherein trailing edge cavity impingement holes (26) are provided in communication between the trailing edge supply cavity (5) and the transverse turbulence column impingement cavity (6).
7. The impingement plenum double wall cooling structure for turbine rotor blades according to claim 1, wherein a third blade back double wall cavity (10) is provided with a third blade back impingement hole (23), and the third blade back impingement hole (23) is communicated with the trailing edge air supply inner cavity (5).
8. The structure of claim 1, wherein the second basin double-wall cavity (12) is provided with a basin trailing edge film hole (18) and a second basin impingement hole (25), the basin trailing edge film hole (18) is communicated with the outside of the rotor blade, and the second basin impingement hole (25) is communicated with the inner trailing edge air supply cavity (5).
9. The double-wall cooling structure for the impact air entrainment film of the turbine rotor blade according to claim 1, characterized in that the transverse turbulence column impact cavity (6), the third blade back double-wall cavity (10) and the second blade basin double-wall cavity (12) are all provided with an outflow seam to communicate with the trailing edge exhaust cavity (7), and the trailing edge exhaust cavity (7) is provided with a trailing edge middle cleft seam (19).
10. An impingement plenum cooling structure for turbine rotor blades according to claim 1 wherein sets of backfoil double-walled cavity turbulators (27) are provided in said third backfoil double-walled cavity (10).
11. An impingement plenum double wall cooling structure for turbine rotor blades according to claim 1 wherein sets of lobe basin double wall cavity turbulator posts (28) are provided within the second lobe basin double wall cavity (12).
12. An impingement plenum double wall cooling structure for turbine rotor blades according to claim 1, wherein sets of transverse inner cavity turbulator posts (29) are provided in the transverse turbulator post impingement cavity (6).
13. The impingement plenum double wall cooling structure for turbine rotor blades according to claim 1, wherein sets of exhaust cavity turbulators (30) are provided within the trailing edge exhaust cavity (7).
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| CN202110526654.8A CN113090335A (en) | 2021-05-14 | 2021-05-14 | Impact air-entraining film double-wall cooling structure for turbine rotor blade |
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| CN114109514A (en) * | 2021-11-12 | 2022-03-01 | 中国航发沈阳发动机研究所 | Turbine blade pressure surface cooling structure |
| CN114151138A (en) * | 2021-10-20 | 2022-03-08 | 中国航发四川燃气涡轮研究院 | Interlayer combined cooling structure of turbine rotor blade |
| CN114575931A (en) * | 2022-03-16 | 2022-06-03 | 中国航发沈阳发动机研究所 | Turbine blade cooling structure with high temperature bearing capacity |
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| CN115095391A (en) * | 2022-06-30 | 2022-09-23 | 上海交通大学 | Turbine blade near-wall surface cooling structure manufactured through additive manufacturing and machining method thereof |
| CN114575931B (en) * | 2022-03-16 | 2024-06-07 | 中国航发沈阳发动机研究所 | Turbine blade cooling structure with high temperature bearing capacity |
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| CN114109514B (en) * | 2021-11-12 | 2023-11-28 | 中国航发沈阳发动机研究所 | Turbine blade pressure surface cooling structure |
| CN114575931A (en) * | 2022-03-16 | 2022-06-03 | 中国航发沈阳发动机研究所 | Turbine blade cooling structure with high temperature bearing capacity |
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| CN114718657A (en) * | 2022-04-08 | 2022-07-08 | 中国航发沈阳发动机研究所 | Local high-efficient cooling structure of turbine blade back of blade |
| CN114718657B (en) * | 2022-04-08 | 2024-06-11 | 中国航发沈阳发动机研究所 | Turbine blade back local high-efficiency cooling structure |
| CN114877375A (en) * | 2022-05-26 | 2022-08-09 | 南京航空航天大学 | Structure for improving double-wall cooling performance by utilizing shape memory alloy |
| CN115095391A (en) * | 2022-06-30 | 2022-09-23 | 上海交通大学 | Turbine blade near-wall surface cooling structure manufactured through additive manufacturing and machining method thereof |
| CN115095391B (en) * | 2022-06-30 | 2023-12-12 | 上海交通大学 | Turbine blade near-wall cooling structure manufactured by additive and processing method thereof |
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