CN113565573A - Turbine blade with internal cooling channels arranged in honeycomb-like manner and gas turbine - Google Patents
Turbine blade with internal cooling channels arranged in honeycomb-like manner and gas turbine Download PDFInfo
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- CN113565573A CN113565573A CN202110769922.9A CN202110769922A CN113565573A CN 113565573 A CN113565573 A CN 113565573A CN 202110769922 A CN202110769922 A CN 202110769922A CN 113565573 A CN113565573 A CN 113565573A
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- turbine blade
- cooling
- honeycomb
- cooling channel
- blade body
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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
- F01D5/147—Construction, i.e. structural features, e.g. of weight-saving hollow 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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- 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)
- Architecture (AREA)
- Turbine Rotor Nozzle Sealing (AREA)
Abstract
The invention provides a turbine blade and a gas turbine with internal cooling channels arranged in a honeycomb-like manner, which comprise: a turbine blade body, cooling passages and film cooling holes; the cooling channel is arranged in the turbine blade body, the film cooling hole is arranged on the surface of the turbine blade body, and the cooling channel is communicated with the film cooling hole; the cooling passage is provided in plurality and divides the inside of the turbine blade body into a honeycomb shape. The efficient utilization of the limited space in the small-size turbine blade is realized, the honeycomb-structure-simulated regular hexagon cooling channel arrangement scheme has the advantages that the length of the cooling channels which can be arranged in the same area is increased, and compared with the square arrangement scheme under the condition of the same channel section equivalent diameter, the limited space in the small-size turbine blade can be more effectively utilized.
Description
Technical Field
The invention relates to the field of design of turbine cooling blades, in particular to a turbine blade with an internal cooling channel arranged in a honeycomb-like manner and a gas turbine.
Background
For a power cycle system in which a turbine set is positioned, increasing the inlet air temperature of a turbine is an important way for improving the cycle efficiency of the system, and the overall cycle efficiency is increased by 20-25% for every 100 ℃ increase of the inlet air temperature of the turbine. For small or micro gas turbine and other equipment, the high temperature working environment exceeding the temperature resistance limit of the material and the small-sized structural design limitation, as well as the requirements of efficient cooling and high stability and long-life operation, all bring higher requirements to the design of the cooling scheme of the turbine blade.
Under high temperature conditions, the guide vanes of the turbine are subject to continuous erosion and corrosion by high temperature gas flow, while the rotor blades are subject to various loads such as centrifugal load, aerodynamic load, vibration load and the like during operation. Particularly in small or micro gas turbine and other equipment, the size of the turbine blade is relatively small, and due to size limitation, the inner cooling channel needs to be a micro-scale inner cooling channel with the equivalent diameter of 0.2-3 mm. In this case, the currently common internal cooling structure of the blade is mostly a straight channel, an S-shaped or tree-shaped branched structure micro-scale channel. For small-size turbine blades which are in service for a long time in a high-temperature environment, an effective cooling channel arrangement scheme is needed, a good cooling effect is achieved in a space with a limited size, and the surface temperature of the turbine blades in the high-temperature environment is reduced to be below the material tolerance limit temperature through an efficient compact cooling structure design; and moreover, the strength of the blades is considered, the maximum stress value of the surfaces of the moving blades with cooling structures in the high-speed rotating process is reduced as much as possible, the safety margin is improved, and the small-size turbine blades can stably run in a long service life under the high-temperature environment.
Patent document CN106555776B relates to a turbofan engine and its fan blades. The fan blade comprises a pressure surface wall plate, a suction surface wall plate and a core plate structure arranged between the pressure surface wall plate and the suction surface wall plate, wherein the core plate structure is provided with a plurality of first joint parts, the positions of the inner side surface of the pressure surface wall plate, which correspond to the first joint parts, are respectively provided with a plurality of bosses, the first joint parts are jointed with the bosses, and the ratio of the height of each boss to the thickness of the core plate structure is 0.5: 1-1: 1.
Disclosure of Invention
In view of the defects in the prior art, the invention aims to provide a turbine blade with an internal cooling channel arranged in a honeycomb-like manner and a gas turbine.
According to the invention, the turbine blade with the inner cooling channels arranged in the honeycomb-like manner comprises: a turbine blade body, cooling passages and film cooling holes;
the cooling channel is arranged in the turbine blade body, the film cooling hole is arranged on the surface of the turbine blade body, and the cooling channel is communicated with the film cooling hole;
the cooling passage is provided in plurality and divides the inside of the turbine blade body into a honeycomb shape.
Preferably, the turbine blade body is provided with a long plate shape, and the tail end of the turbine blade body is provided with a blade bottom mortise;
preferably, one end of the turbine blade body, which faces away from the blade bottom mortise, is set as a top, one side of the turbine blade body is set as a leading edge, and the other side of the turbine blade body is set as a trailing edge.
Preferably, the film cooling holes are provided at the top, at the leading edge and at the trailing edge.
Preferably, one end of the turbine blade body, which is provided with the blade bottom mortise, is provided with an air inlet;
the air inlet is communicated with the cooling channel.
Preferably, when a cooling air flow flows from the air inlet to the film cooling hole through the cooling passage, the cooling air flow passes through the film cooling hole and forms a cooling film.
Preferably, the cooling channels are connected end to form a regular hexagon unit, and the interior of the turbine blade body is divided into a honeycomb shape by the regular hexagon unit and is provided with a plurality of regular hexagon columns.
Preferably, the regular hexagonal pillars are in heat exchange with the cooling airflow.
Preferably, the cooling channel is rectangular in cross section and the air inlet is rectangular in cross section.
Preferably, the internal cooling channels of the gas turbine are arranged in a honeycomb-like manner.
Preferably, the equivalent diameter of the cooling channel cross-section is 0.2-3 mm.
Preferably, the film cooling hole is cylindrical, and the diameter of the film cooling hole is smaller than or equal to 1 mm.
Preferably, the equivalent diameter of the section of the air inlet is 4 to 5 times that of the section of the cooling channel.
Preferably, the turbine blade body is a small-sized turbine blade, the thickness of the turbine blade body is less than 10mm, and the chord length of the turbine blade body is less than 30 mm;
the thickness is the turbine blade body height along the regular hexagon post axial direction, and the chord length is the turbine blade body length along the direction that extends at the leading edge.
Compared with the prior art, the invention has the following beneficial effects:
1. the efficient utilization of the limited space in the small-size turbine blade is realized, the honeycomb-structure-simulated regular hexagon cooling channel arrangement scheme has the advantages that the length of the cooling channels which can be arranged in the same area is increased, and compared with the square arrangement scheme under the condition of the same channel section equivalent diameter, the limited space in the small-size turbine blade can be more effectively utilized.
2. The invention combines the cooling of the micro-scale channels inside the blades and the cooling of the external air film, and the cooling air is continuously polymerized and distributed in the micro-scale cooling channels which are arranged in the internal honeycomb-like structure, thereby greatly improving the internal flow turbulence degree and enhancing the heat exchange efficiency, and further achieving the purpose of efficiently cooling the small-size turbine blades.
3. The strength of the small-size turbine blade with the cooling structure is guaranteed, the cooling structure inside the blade is formed by communicating a plurality of adjacent arranged regular hexagon micro-channels, the whole distribution is honeycomb-shaped, the structural strength of the hollow blade can be well guaranteed by the arrangement mode, the capability of resisting various loads such as centrifugal load, pneumatic load and vibration load in multiple directions is realized, and therefore the strength safety margin and service life of the blade material are increased.
Drawings
Other features, objects and advantages of the invention will become more apparent upon reading of the detailed description of non-limiting embodiments with reference to the following drawings:
FIG. 1 is a schematic structural view of a turbine blade with internal cooling channels arranged in a honeycomb-like manner;
shown in the figure:
Detailed Description
The present invention will be described in detail with reference to specific examples. The following examples will assist those skilled in the art in further understanding the invention, but are not intended to limit the invention in any way. It should be noted that it would be obvious to those skilled in the art that various changes and modifications can be made without departing from the spirit of the invention. All falling within the scope of the present invention.
Example 1
As shown in fig. 1, a turbine blade having an internal cooling channel arranged in a honeycomb-like manner includes: a turbine blade body 1, a cooling passage 4, and a film cooling hole; inside cooling channel 4 that sets up of turbine blade body 1, turbine blade body 1 surface sets up the air film cooling hole, and cooling channel 4 intercommunication air film cooling hole, cooling channel 4 are provided with a plurality ofly and are divided into honeycombedly with turbine blade body 1 is inside. The turbine blade body 1 is in a long plate shape, and the tail end of the turbine blade body 1 is provided with a blade bottom mortise 2; one end of the turbine blade body 1, which faces away from the blade bottom mortise 2, is provided as a top 6, one side of the turbine blade body 1 is provided as a leading edge 5, and the other side of the turbine blade body 1 is provided as a trailing edge 7. Top 6, leading edge department 5 and trailing edge department 7 set up the film cooling hole, and turbine blade body 1 sets up 2 one ends of blade bottom tongue-and-groove and sets up air inlet 3, and air inlet 3 intercommunication cooling channel 4, and the cooling channel 4 cross-section sets up to the rectangle, and the 3 cross-sections of air inlet set up to the rectangle.
A plurality of cooling channels 4 end to end connect for regular hexagon unit, and turbine blade body 1 is inside to be cellular and set up a plurality of regular hexagon posts through a plurality of regular hexagon unit divisions. When the cooling air flow flows from the air inlet 3 to the film cooling holes through the cooling passages 4, the cooling air flow passes through the film cooling holes and forms a cooling film, and the regular hexagonal columns exchange heat with the cooling air flow. Gas turbines may employ the present invention.
Example 2
The turbine blade with the internal cooling channels arranged in the honeycomb-like mode comprises a turbine blade body 1 and a blade bottom mortise 2, wherein the turbine blade body 1 is internally provided with micro-scale cooling channels 4 arranged according to a honeycomb structure, and film cooling holes which are formed in the surface of the turbine blade body 1 at the front edge 5, the top 6 and the tail edge 7 of the blade are communicated with the internal cooling channels 4. And an air inlet 3 for cooling air flow is arranged at the mortise 2 at the bottom of the blade. The blade thickness of the turbine blade body 1 is less than 10mm, the chord length of the turbine blade body 1 is less than 30mm, and the turbine blade is a small-size turbine blade suitable for a small-size or micro gas turbine. The section of the micro-scale cooling channel 4 which is arranged in the honeycomb-like structure in the turbine blade body 1 is rectangular, the equivalent diameter is 0.2-3mm, the turbine blade is formed by communicating a plurality of adjacent regular hexagonal micro-channels, and the whole distribution is honeycomb-shaped. Cooling airflow enters a cooling channel 4 which is arranged in a honeycomb-like structure in the blade from an air inlet 3 at a mortise 2 at the bottom of the blade, the section of the air inlet 3 is rectangular, and the equivalent diameter of the air inlet is 4-5 times that of the internal main cooling channel 4. The cooling air flow is finally discharged through film cooling holes distributed at the leading edge 5, the top 6 and the trailing edge 7, the film cooling holes are cylindrical air flow channels with the diameter smaller than 1mm and are communicated with the outer surface of the blade and the inner main cooling channel 4. The cooling airflow enters the cooling channel 4 from the air inlet 3 at the mortise 2 at the bottom of the blade, and exchanges heat with the wall surface of the regular hexagon column in the high-temperature blade in the cooling channel 4 in a flowing mode, so that the surface temperature of the blade is reduced, and finally the cooling airflow is discharged through film cooling holes distributed at the front edge 5, the top 6 and the tail edge 7 of the blade to form a layer of cooling film, and the blade is further cooled.
The invention realizes the efficient utilization of the limited space in the small-size turbine blade. The honeycomb-like regular hexagon cooling channel 4 arrangement scheme can increase the length of the cooling channel 4 which can be arranged in the same area by more than 10% compared with the square arrangement scheme under the condition of the same channel section equivalent diameter, and can more effectively utilize the limited space in the small-size turbine blade. According to the invention, the cooling of the micro-scale channel inside the blade and the cooling of the external air film are combined, and through simulation evaluation, the comprehensive cooling effect can reduce the average temperature of the surface of the small-size turbine blade with the blade height of 8mm and the blade chord length of 18mm by about 200K. The cooling gas is continuously polymerized and distributed in the small-scale cooling channel arranged in the internal honeycomb-like structure, so that the internal flow turbulence is greatly improved, the heat exchange efficiency is enhanced, and the purpose of efficiently cooling the small-size turbine blades is achieved. The invention ensures the strength of the small-size turbine blade with the cooling structure, the cooling structure in the blade is formed by communicating a plurality of adjacent regular hexagonal micro-channels, the whole distribution is in a honeycomb shape, the arrangement mode can well ensure the structural strength of the hollow blade, and the capability of resisting various loads such as centrifugal load, pneumatic load, vibration load and the like in multiple directions is realized. Through simulation evaluation, compared with the common blade with the straight-through type internal cooling channel, the blade provided by the invention has the advantages that the maximum stress value of the surface of the moving blade can be reduced by more than 10% in the high-speed working process, so that the strength safety margin and the service life of the blade material are increased.
Specifically, the thickness of the small-size turbine blade with the inner cooling channels arranged in a honeycomb-like manner is 9mm, the chord length of the blade is 18mm, and the cooling structure is used for a small gas turbine. The section of the cooling channel 4 is rectangular, the equivalent diameter is 1mm, the cooling channel is formed by communicating a plurality of adjacent regular hexagonal micro-channels, and the cooling channel is integrally distributed in a honeycomb shape. The cross section of the air inlet 3 is rectangular, and the equivalent diameter is 5mm which is 5 times of the equivalent diameter of the inner main cooling channel. The film cooling holes arranged at the leading edge 5, the tip 6 and the trailing edge 7 of the blade are cylindrical channels of air flow with a diameter of 0.3mm, communicating the outer surface of the blade with the inner main cooling channel 4. The cooling air flow enters the cooling channel 4 which is arranged in the blade in a honeycomb-like structure from the air inlet 3 at the mortise 2 at the bottom of the blade, and carries out heat convection with the inner wall surface of the high-temperature blade in the cooling channel 4, so that the surface temperature of the blade is reduced, and finally the cooling air flow is discharged through air film cooling holes which are distributed at the front edge 5, the top 6 and the tail edge 7 of the blade to form a layer of cooling air film, and the blade is further cooled.
In the description of the present application, it is to be understood that the terms "upper", "lower", "front", "rear", "left", "right", "vertical", "horizontal", "top", "bottom", "inner", "outer", and the like indicate orientations or positional relationships based on those shown in the drawings, and are only for convenience in describing the present application and simplifying the description, but do not indicate or imply that the referred device or element must have a specific orientation, be constructed in a specific orientation, and be operated, and thus, should not be construed as limiting the present application.
The foregoing description of specific embodiments of the present invention has been presented. It is to be understood that the present invention is not limited to the specific embodiments described above, and that various changes or modifications may be made by one skilled in the art within the scope of the appended claims without departing from the spirit of the invention. The embodiments and features of the embodiments of the present application may be combined with each other arbitrarily without conflict.
Claims (10)
1. A turbine blade with internal cooling channels arranged in a honeycomb-like manner, comprising: the turbine blade comprises a turbine blade body (1), a cooling channel (4) and an air film cooling hole;
the cooling channel (4) is arranged in the turbine blade body (1), the film cooling hole is arranged on the surface of the turbine blade body (1), and the cooling channel (4) is communicated with the film cooling hole;
the cooling channel (4) is provided with a plurality of cooling channels and divides the interior of the turbine blade body (1) into a honeycomb shape.
2. The internal cooling channel honeycomb-like arrangement turbine blade of claim 1, wherein: the turbine blade body (1) is arranged to be long plate-shaped, and a blade bottom mortise (2) is arranged at the tail end of the turbine blade body (1).
3. The internal cooling channel honeycomb-like arrangement turbine blade of claim 2, wherein: turbine blade body (1) dorsad blade bottom tongue-and-groove (2) one end sets up to top (6), turbine blade body (1) one side sets up to leading edge department (5), turbine blade body (1) another side sets up to trailing edge department (7).
4. The internal cooling channel honeycomb-like arrangement turbine blade of claim 3, wherein: the top (6), the leading edge (5) and the trailing edge (7) are provided with the film cooling holes.
5. The internal cooling channel honeycomb-like arrangement turbine blade of claim 4, wherein: one end of the turbine blade body (1) provided with the blade bottom mortise (2) is provided with an air inlet (3);
the air inlet (3) is communicated with the cooling channel (4).
6. The internal cooling channel honeycomb-like arrangement turbine blade of claim 5, wherein: when a cooling air flow flows from the air inlet (3) to the film cooling hole through the cooling passage (4), the cooling air flow passes through the film cooling hole and forms a cooling film.
7. The internal cooling channel honeycomb-like arrangement turbine blade of claim 6, wherein: a plurality of cooling channel (4) end to end connect is regular hexagon unit, turbine blade body (1) is inside through a plurality of regular hexagon unit is cut apart into honeycomb and is set up a plurality of regular hexagon post.
8. The internal cooling channel honeycomb-like arrangement turbine blade of claim 7, wherein: the regular hexagonal pillars achieve heat exchange with the cooling airflow.
9. The internal cooling channel honeycomb-like arrangement turbine blade of claim 5, wherein: the section of the cooling channel (4) is rectangular;
the cross section of the air inlet (3) is rectangular.
10. A gas turbine, characterized by: a gas turbine employing a turbine blade having an internal cooling channel according to any one of claims 1 to 9 in a honeycomb-like arrangement.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202110769922.9A CN113565573B (en) | 2021-07-07 | 2021-07-07 | Turbine blade with internal cooling channels distributed in honeycomb-like manner and gas turbine |
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| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202110769922.9A CN113565573B (en) | 2021-07-07 | 2021-07-07 | Turbine blade with internal cooling channels distributed in honeycomb-like manner and gas turbine |
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| CN113565573A true CN113565573A (en) | 2021-10-29 |
| CN113565573B CN113565573B (en) | 2023-08-11 |
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Cited By (2)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN114863783A (en) * | 2022-05-27 | 2022-08-05 | 中国科学院工程热物理研究所 | Turbine blade leading edge simulation piece |
| CN114961895A (en) * | 2022-06-16 | 2022-08-30 | 大连理工大学 | Turbine outer ring adopting double-helix cooling structure |
Citations (8)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US6235370B1 (en) * | 1999-03-03 | 2001-05-22 | Siemens Westinghouse Power Corporation | High temperature erosion resistant, abradable thermal barrier composite coating |
| JP2002174102A (en) * | 2000-12-07 | 2002-06-21 | Ishikawajima Harima Heavy Ind Co Ltd | Transpiration cooling heat transfer promotion structure of turbine blade |
| US20100254806A1 (en) * | 2009-04-06 | 2010-10-07 | General Electric Company | Methods, systems and/or apparatus relating to seals for turbine engines |
| CN106555776A (en) * | 2015-09-25 | 2017-04-05 | 中航商用航空发动机有限责任公司 | Turbofan and its fan blade |
| CN111075510A (en) * | 2020-01-06 | 2020-04-28 | 大连理工大学 | Turbine blade honeycomb spiral cavity cooling structure |
| CN111140287A (en) * | 2020-01-06 | 2020-05-12 | 大连理工大学 | Laminate cooling structure adopting polygonal turbulence column |
| CN111305906A (en) * | 2020-03-31 | 2020-06-19 | 哈尔滨工程大学 | Area is disconnected straight rib and is half split joint cooling structure between suitable for high temperature turbine blade |
| CN112145234A (en) * | 2020-09-24 | 2020-12-29 | 大连理工大学 | Omega type gyration chamber plywood cooling structure |
-
2021
- 2021-07-07 CN CN202110769922.9A patent/CN113565573B/en active Active
Patent Citations (8)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US6235370B1 (en) * | 1999-03-03 | 2001-05-22 | Siemens Westinghouse Power Corporation | High temperature erosion resistant, abradable thermal barrier composite coating |
| JP2002174102A (en) * | 2000-12-07 | 2002-06-21 | Ishikawajima Harima Heavy Ind Co Ltd | Transpiration cooling heat transfer promotion structure of turbine blade |
| US20100254806A1 (en) * | 2009-04-06 | 2010-10-07 | General Electric Company | Methods, systems and/or apparatus relating to seals for turbine engines |
| CN106555776A (en) * | 2015-09-25 | 2017-04-05 | 中航商用航空发动机有限责任公司 | Turbofan and its fan blade |
| CN111075510A (en) * | 2020-01-06 | 2020-04-28 | 大连理工大学 | Turbine blade honeycomb spiral cavity cooling structure |
| CN111140287A (en) * | 2020-01-06 | 2020-05-12 | 大连理工大学 | Laminate cooling structure adopting polygonal turbulence column |
| CN111305906A (en) * | 2020-03-31 | 2020-06-19 | 哈尔滨工程大学 | Area is disconnected straight rib and is half split joint cooling structure between suitable for high temperature turbine blade |
| CN112145234A (en) * | 2020-09-24 | 2020-12-29 | 大连理工大学 | Omega type gyration chamber plywood cooling structure |
Cited By (3)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN114863783A (en) * | 2022-05-27 | 2022-08-05 | 中国科学院工程热物理研究所 | Turbine blade leading edge simulation piece |
| CN114863783B (en) * | 2022-05-27 | 2024-02-27 | 中国科学院工程热物理研究所 | Turbine blade leading edge simulation piece |
| CN114961895A (en) * | 2022-06-16 | 2022-08-30 | 大连理工大学 | Turbine outer ring adopting double-helix cooling structure |
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| CN113565573B (en) | 2023-08-11 |
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