CN113944516A - Gas turbine blade tip composite cooling structure - Google Patents
Gas turbine blade tip composite cooling structure Download PDFInfo
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- CN113944516A CN113944516A CN202111142033.6A CN202111142033A CN113944516A CN 113944516 A CN113944516 A CN 113944516A CN 202111142033 A CN202111142033 A CN 202111142033A CN 113944516 A CN113944516 A CN 113944516A
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- gas turbine
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- 238000001816 cooling Methods 0.000 title claims abstract description 110
- 239000002131 composite material Substances 0.000 title claims abstract description 30
- 239000012530 fluid Substances 0.000 claims abstract description 59
- 230000009471 action Effects 0.000 claims abstract description 10
- 238000009991 scouring Methods 0.000 claims abstract description 5
- 239000007789 gas Substances 0.000 claims description 98
- 238000010408 sweeping Methods 0.000 claims description 17
- 239000002737 fuel gas Substances 0.000 claims description 3
- 238000005507 spraying Methods 0.000 claims description 3
- 239000004047 hole gas Substances 0.000 claims 1
- 230000009286 beneficial effect Effects 0.000 abstract description 4
- 238000009827 uniform distribution Methods 0.000 abstract description 2
- 238000009826 distribution Methods 0.000 description 8
- 238000007664 blowing Methods 0.000 description 4
- 230000000694 effects Effects 0.000 description 4
- 238000013461 design Methods 0.000 description 3
- 238000010586 diagram Methods 0.000 description 3
- 150000001875 compounds Chemical class 0.000 description 2
- 238000007789 sealing Methods 0.000 description 2
- 239000000463 material Substances 0.000 description 1
- 238000002844 melting Methods 0.000 description 1
- 230000008018 melting Effects 0.000 description 1
- 238000000034 method Methods 0.000 description 1
- 238000012986 modification Methods 0.000 description 1
- 230000004048 modification Effects 0.000 description 1
- 230000000737 periodic effect Effects 0.000 description 1
- 230000008569 process Effects 0.000 description 1
- 238000012545 processing Methods 0.000 description 1
- 239000007787 solid Substances 0.000 description 1
- 230000003068 static effect Effects 0.000 description 1
- 238000012546 transfer Methods 0.000 description 1
- 238000013519 translation Methods 0.000 description 1
Images
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
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- F05D2260/00—Function
- F05D2260/20—Heat transfer, e.g. cooling
- F05D2260/201—Heat transfer, e.g. cooling by impingement of a fluid
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- F05D2260/00—Function
- F05D2260/20—Heat transfer, e.g. cooling
- F05D2260/202—Heat transfer, e.g. cooling by film cooling
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- Engineering & Computer Science (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Turbine Rotor Nozzle Sealing (AREA)
Abstract
The invention discloses a gas turbine blade tip composite cooling structure combining sweep impact and a gas film hole, wherein an impact cooling cavity and a high-temperature gas main flow cavity are arranged at two sides of a gas turbine blade tip flat plate model, an outlet section of a fluid vibration exciter is arranged in the impact cooling cavity and is opposite to the inner surface of the flat plate model, a cooling working medium is sprayed from the outlet section of the fluid vibration exciter into the impact cooling cavity to form sweep impact jet flow acting on an impact target surface, the turbulent kinetic energy of the cooling working medium near the impact target surface is increased under the suction action of the gas film hole, and the cooling working medium entering the high-temperature gas main flow cavity through the gas film hole forms a gas film layer covering the outer surface of the flat plate model so as to avoid direct scouring of the high-temperature main flow gas on the outer surface of the gas turbine blade tip flat plate model. The composite cooling structure is beneficial to avoiding local hot spots at the tip of the gas turbine blade, and has the advantages of large cooling area, uniform distribution of cooling efficiency, small cold air consumption ratio and the like.
Description
Technical Field
The invention belongs to the technical field of turbine blade cooling, relates to a gas turbine blade tip composite cooling structure, and particularly relates to a gas turbine blade tip composite cooling structure combining sweeping impact and a gas film hole, which is beneficial to avoiding local hot spots of a gas turbine blade tip, has the characteristics of simple structure, high cooling performance, less amount of adopted cold gas and the like, and is particularly suitable for a high-heat-load gas-cooled turbine stator blade front edge.
Background
According to the thermodynamic law in gas turbines, the brayton cycle shows that the operating efficiency of a gas turbine increases with increasing turbine inlet temperature. At present, the front total temperature of the turbine reaches about 2200K, and the melting point of the blade material of the modern gas turbine is only about 1500K, so that the cooling design of the modern gas turbine needs to meet the technical requirements of high cooling efficiency, low aerodynamic loss, small air-cooling air volume ratio and the like in order to continuously improve the working efficiency and avoid the thermal deformation of the blade. The blade tip of the first stage turbine guider is firstly subjected to high-temperature gas scouring, so that the arrangement of a cooling structure which can improve the cooling efficiency of the region and avoid hot spots and the like is one of important contents of the cooling design of the gas turbine.
At present, the cooling structure of the turbine blade tip mainly comprises direct impact and air film hole composite cooling or vortex impact and air film hole composite cooling, and compared with the direct impact and air film hole composite cooling, the cooling efficiency of the vortex impact and air film hole composite cooling is more uniform in distribution, but the cooling efficiency of a vortex impact core area is lower than that of the direct impact and air film hole composite cooling. Due to the sweeping action of the fluid vibration exciter, the sweeping impact and the air film hole compound cooling not only enable the cooling efficiency to be uniformly distributed, but also can save cold air quantity, so that the adoption of the sweeping impact and air film hole compound cooling is beneficial to avoiding the occurrence of local hot spots on the front edge and reducing the cold air quantity ratio.
Disclosure of Invention
In view of the above disadvantages and shortcomings of the prior art, an object of the present invention is to provide a composite cooling structure for a gas turbine blade tip combining sweep impact with a film hole, which is beneficial to avoiding local hot spots on the gas turbine blade tip, and has the advantages of large cooling area, uniform distribution of cooling efficiency, small air-cooling usage ratio, and the like, compared with the existing composite cooling structure for a blade tip combining static impact such as direct impact or vortex impact with a film hole.
In order to solve the technical problem, the invention adopts the technical scheme that:
a gas turbine blade tip composite cooling structure combining sweeping impact and gas film holes comprises a fluid vibration exciter, an impact cooling cavity, a high-temperature gas main flow cavity and a gas turbine blade tip flat plate model, wherein the gas film holes are arranged in the gas turbine blade tip flat plate model in a sequential manner, the fluid vibration exciter comprises an inlet section, a mixing cavity, a loop and an outlet section, and is characterized in that,
the impingement cooling cavity is arranged on one side of the gas turbine blade tip flat plate model, the high-temperature gas main flow cavity is arranged on the other side of the gas turbine blade tip flat plate model,
at least the outlet section of the fluid vibration exciter is arranged in the impact cooling cavity, the outlet section of the fluid vibration exciter is right opposite to the inner surface of the gas turbine blade tip flat model, the inner surface of the gas turbine blade tip flat model is formed into an impact target surface, a cooling working medium is sprayed into the impact cooling cavity from the outlet section of the fluid vibration exciter, a sweeping impact jet is formed under the high-frequency sweeping action of the fluid vibration exciter and acts on the impact target surface, the turbulent kinetic energy of the cooling working medium near the impact target surface is increased under the suction action of the gas film hole, the cooling working medium entering the high-temperature gas main flow cavity through the gas film hole is formed into a gas film layer covering the outer surface of the gas turbine blade tip flat model, so as to avoid the direct scouring of the high-temperature main flow gas on the outer surface of the gas turbine blade tip flat model, and the cooling working medium after heat exchange is discharged through an outlet of the high-temperature gas main flow cavity.
In the gas turbine blade tip composite cooling structure combining the sweep impact and the gas film hole, the cooling working medium is sprayed from the fluid vibration exciter into the impact cooling cavity, and the sweep impact jet flow is formed under the high-frequency sweeping action of the fluid vibration exciter to act on the impact target surface.
Preferably, the impingement cooling chamber has a height H for a desired heat exchange effect1The ratio H to the hydraulic diameter D of the throat of the fluid vibration exciter1D is 1-15, and the height H of the high-temperature fuel gas main flow cavity is used for meeting the requirements of blade tip gaps with different sizes2The ratio H to the hydraulic diameter D of the throat of the fluid vibration exciter2the/D is 2-10, and in order to achieve an ideal heat exchange effect, the dimensionless spraying distance S/D from the throat part of the fluid vibration exciter to the impact target surface is 2-16.
Preferably, in order to produce a suitable tip radial cooling area, the fluid vibration exciter is a rounded corner type vibration exciter, the throat aspect ratio of the fluid vibration exciter is 0.5-3, and in order to produce a suitable tip axial cooling area, the ratio h/D of the height h of the outlet section to the throat hydraulic diameter D is 0.5-3.
Preferably, in order to meet working conditions of different cold air volumes, the ratio D/D of the diameter D of each air film hole to the hydraulic diameter D of the throat part of the fluid vibration exciter is 0.5-3, in order to achieve different cooling efficiencies, the ratio of the distance between adjacent rows of air film holes to the hydraulic diameter D of the throat part of the fluid vibration exciter is 0.5-20, the ratio of the distance between adjacent air exhaust film holes to the hydraulic diameter D of the throat part of the fluid vibration exciter is 0.5-20, and in order to give consideration to the requirements of air film cooling efficiency and blade tip sealing, the air film holes are inclined along the main flow direction, and the angle is 10-90 degrees.
Preferably, for being applied to turbine blades with different chord lengths, the ratio L/D of the length L of the gas turbine blade tip flat plate model to the hydraulic diameter D of the throat part of the fluid vibration exciter is 1.5-200, for being applied to turbine blades with different thicknesses, the ratio W/D of the width W to the hydraulic diameter D of the throat part of the fluid vibration exciter is 0.5-100, for being applied to turbine blades with different blade tip thicknesses, the thickness H of the turbine blade is3The ratio H of the hydraulic diameter D of the throat part of the fluid vibration exciter3the/D is 0.5-10.
Compared with the prior art, the gas turbine blade tip composite cooling structure combining the sweeping impact and the gas film hole mainly has the following technical effects:
1. the cooling zones are different: the composite cooling structure combining the sweeping impact and the air film hole is suitable for the tip area of the gas turbine with the flat impact target surface. Meanwhile, the invention standardizes the value range of each geometric parameter, so that the application range of the invention is more definite, and the heat exchange efficiency is controlled in an effective range.
2. The structure is simple: the design similar to the traditional direct impact and air film composite cooling is adopted, the reliability is higher, and the processing is convenient.
3. The cooling performance is high: compared with direct impact and air film composite cooling, the cooling area of sweep impact and air film composite cooling is larger, and the cooling efficiency distribution is more uniform.
4. The amount of cold air adopted is less: because the cooling effect of the direct impact and air film cooling composite structure can be achieved by adopting less cold air, the use amount of the cold air of the turbine blade is saved, and the turbine efficiency is further improved.
Drawings
FIG. 1 is a schematic view of the present invention of a gas turbine tip composite cooling configuration incorporating swept impingement and film holes.
Fig. 2 is a schematic structural diagram of a fluid vibration exciter.
FIG. 3 is a schematic view of a gas turbine blade tip flat plate model.
In fig. 4, (a) is a schematic diagram showing the influence of the blowing ratio on the distribution of the Nu numbers of the swept and direct impact surfaces; (b) the effect of the blowing ratio on the cooling efficiency η distribution of the swept and direct impingement plate outer surface is illustrated schematically.
FIG. 5 is a schematic view of a centerline integrated film cooling efficiency curve for the outer surface of a gas turbine bucket tip plate model.
Description of reference numerals:
the device comprises a fluid vibration exciter 1, an impact cooling cavity 2, a gas film hole 3, a high-temperature gas main flow cavity 4, a gas turbine blade tip flat plate model 5 and a fluid flowing direction 6.
Detailed Description
In order to make the implementation objects, technical solutions and advantages of the present invention clearer, the technical solutions in the embodiments of the present invention will be described in more detail below with reference to the accompanying drawings in the embodiments of the present invention. In the drawings, the same or similar reference numerals denote the same or similar elements or elements having the same or similar functions throughout. The described embodiments, which are part of the present invention, are not all embodiments, and are intended to be illustrative of the present invention and should not be construed as limiting the present invention. 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.
Fig. 1 shows a gas turbine blade tip composite cooling structure combining sweep impact and a gas film hole, which is provided by the invention, and the gas turbine blade tip composite cooling structure comprises a fluid vibration exciter 1, an impact cooling cavity 2, a gas film hole 3, a high-temperature gas mainstream cavity 4 and a gas turbine blade tip flat plate model 5. The impingement cooling cavity 2 and the high-temperature fuel gas main flow cavity 4 are cuboids; the fluid vibration exciter 1 is arranged above the impingement cooling cavity 2; the gas turbine blade tip flat plate model 5 is arranged between the impingement cooling cavity 2 and the high-temperature gas main flow cavity 4; the film holes 3 are arranged on the gas turbine blade tip flat plate model 5 in a row. At least the outlet section of the fluid vibration exciter 1 is arranged in the impingement cooling cavity 2, and the outlet section of the fluid vibration exciter 1 is facing the inner surface of the gas turbine blade tip flat model 5, and the inner surface of the gas turbine blade tip flat model 5 is formed into an impingement target surface.
In the gas turbine blade tip composite cooling structure combining the sweep impact and the gas film hole, the whole flowing heat transfer process is as follows: the cooling working medium is sprayed from the outlet section of the fluid vibration exciter 1 to enter the impact cooling cavity 2, a sweeping impact jet flow is formed under the action of a high-frequency sweeping action and acts on an impact target surface, compared with direct impact, the coverage area of the cooling working medium on the impact target surface is larger, heat exchange is more uniform, turbulent kinetic energy of the cooling working medium nearby the impact target surface is increased under the suction action of the air film hole 3, the formed air film covers the outer surface of the gas turbine blade tip flat plate model 5, direct scouring of high-temperature gas is avoided, the cooling working medium after heat exchange is discharged through the outlet of the high-temperature gas main flow cavity 4, and the fluid flowing direction 6 of the cooling working medium in the high-temperature gas main flow cavity 4 is shown as an arrow in figure 1.
Referring to fig. 1, the impingement cooling chamber 2 is a rectangular parallelepiped, and has a length and width equal to those of the gas turbine blade tip flat mold 5 and a height H for achieving an ideal heat exchange effect1The ratio H of the hydraulic diameter D of the throat part of the fluid vibration exciter 11and/D is 1-15, so that the dimensionless spraying distance S/D from the throat part of the fluid vibration exciter 1 to the impact target surface is 2-16. The high-temperature gas main flow cavity 4 is a cuboid, the length and the width of the high-temperature gas main flow cavity are the same as those of the flat plate model 5, and the height H of the high-temperature gas main flow cavity is equal to that of the blade top clearance of different sizes2The ratio H of the hydraulic diameter D of the throat part of the fluid vibration exciter 12The value of/D is 2-10. In order to ensure that the main flow boundary layer is fully developed during numerical calculation, the main flow cavity 4 extends forwards by half of the length of the main flow cavity. Both sides of the impingement cooling cavity 2, the high temperature gas mainstream cavity 4 and the gas turbine blade tip flat model 5 are set as periodic translation boundaries, and thus, the above is only one cooling unit of the infinite long flat model.
The fluid vibration exciter shown in fig. 2 is of a round angle type and comprises an inlet section, a mixing cavity, a loop and an outlet section, wherein the throat aspect ratio of the fluid vibration exciter is 0.5-3 for generating a proper blade tip radial cooling area, and the ratio h/D of the height h of the outlet section to the throat hydraulic diameter D is 0.5-3 for generating a proper blade tip axial cooling area.
Fig. 3 is a schematic diagram of a gas turbine blade tip flat plate model, in order to meet working conditions of different amounts of cold air, the ratio D/D of the diameter of a gas film hole to the hydraulic diameter of the throat of a fluid vibration exciter is 0.5-3, in order to achieve different cooling efficiencies, the ratio of the distance between adjacent rows of gas film holes to the hydraulic diameter of the throat of the fluid vibration exciter is 0.5-20, the ratio of the distance between adjacent exhaust film holes to the hydraulic diameter of the throat of the fluid vibration exciter is 0.5-20, and in order to give consideration to the requirements of gas film cooling efficiency and blade tip sealing, the gas film holes are inclined along the main flow direction, and the angle is 10-90 degrees. In addition, for being applied to turbine blades with different chord lengths, the L/D ratio of the length of the flat plate model to the hydraulic diameter of the throat part of the fluid vibration exciter is 1.5-200, for being applied to turbine blades with different thicknesses, the W/D ratio of the width to the hydraulic diameter of the throat part of the fluid vibration exciter is 0.5-100, for being applied to turbine blades with different blade tip thicknesses, the H/D ratio of the height to the hydraulic diameter of the throat part of the fluid vibration exciter3The ratio of/D is 0.5-10, in order to make the inside of the solid domainThe pythagorean number of (a) is similar to that of a real gas turbine, and the thermal conductivity thereof is set to 5-15W/(s.K).
Fig. 4 shows the distribution of the number Nu of the inner surfaces of the flat plates and the distribution of the cooling efficiency (expressed by η) of the outer surfaces when the blowing ratio (expressed by BR) is 1, and fig. 5 shows the comprehensive film cooling efficiency curve of the center line of the outer surfaces of the flat plate models of the tips of the gas turbines. Compared with direct impact (expressed as DJ), the cold air effective coverage area of the sweep impact (expressed as SJ) is larger (the area of the sweep impact surface Nu number is more than 20 is 1.91 times of that of the direct impact on average) and is more uniform. And when the dimensionless impact distance is equal to 5, the area average cooling efficiency of the sweep impact is 22.49% -26.00% higher, and the increase speed of the cooling efficiency of the sweep impact is increased along with the increase of the blowing ratio. Therefore, although the average flow resistance coefficient is 2 times of that of the direct impact, the cold air flow distribution capacity of the fluid vibration exciter is high, and the distribution area and the numerical value of the comprehensive cooling efficiency are relatively high.
The object of the present invention is fully effectively achieved by the above embodiments. Those skilled in the art will appreciate that the present invention includes, but is not limited to, what is described in the accompanying drawings and the foregoing detailed description. While the invention has been described in connection with what is presently considered to be the most practical and preferred embodiment, it is to be understood that the invention is not to be limited to the disclosed embodiment, but on the contrary, is intended to cover various modifications within the spirit and scope of the appended claims.
Claims (5)
1. A gas turbine blade tip composite cooling structure combining sweeping impact and gas film holes comprises a fluid vibration exciter, an impact cooling cavity, a high-temperature gas main flow cavity and a gas turbine blade tip flat plate model, wherein the gas film holes are arranged in the gas turbine blade tip flat plate model in a sequential manner, the fluid vibration exciter comprises an inlet section, a mixing cavity, a loop and an outlet section, and is characterized in that,
the impingement cooling cavity is arranged on one side of the gas turbine blade tip flat plate model, the high-temperature gas main flow cavity is arranged on the other side of the gas turbine blade tip flat plate model,
at least the outlet section of the fluid vibration exciter is arranged in the impact cooling cavity, the outlet section of the fluid vibration exciter is right opposite to the inner surface of the gas turbine blade tip flat model, the inner surface of the gas turbine blade tip flat model is formed into an impact target surface, a cooling working medium is sprayed into the impact cooling cavity from the outlet section of the fluid vibration exciter, a sweeping impact jet is formed under the high-frequency sweeping action of the fluid vibration exciter and acts on the impact target surface, the turbulent kinetic energy of the cooling working medium near the impact target surface is increased under the suction action of the gas film hole, the cooling working medium entering the high-temperature gas main flow cavity through the gas film hole is formed into a gas film layer covering the outer surface of the gas turbine blade tip flat model, so as to avoid the direct scouring of the high-temperature main flow gas on the outer surface of the gas turbine blade tip flat model, and the cooling working medium after heat exchange is discharged through an outlet of the high-temperature gas main flow cavity.
2. The combined swept impingement and film hole gas turbine blade tip composite cooling structure of the preceding claims, wherein the impingement cooling cavity height H1The ratio H to the hydraulic diameter D of the throat of the fluid vibration exciter1D is 1-15, and the height H of the high-temperature fuel gas main flow cavity2The ratio H to the hydraulic diameter D of the throat of the fluid vibration exciter2And the dimensionless spraying distance S/D from the throat part of the fluid vibration exciter to the impact target surface is 2-16.
3. The gas turbine blade tip composite cooling structure combining sweeping impact and film hole as claimed in the preceding claim, wherein the fluid exciter is a rounded corner exciter with a throat aspect ratio of 0.5-3 and a ratio h/D of a height h of the outlet section to a throat hydraulic diameter D of 0.5-3.
4. The composite cooling structure of the gas turbine blade tip combining the sweep impact and the air film hole as claimed in the previous claims, wherein the ratio D/D of the diameter D of the air film hole to the hydraulic diameter D of the throat of the fluid exciter is 0.5-3, the ratio of the distance between the air film holes of adjacent rows to the hydraulic diameter D of the throat of the fluid exciter is 0.5-20, the ratio of the distance between the adjacent air discharge film holes to the hydraulic diameter D of the throat of the fluid exciter is 0.5-20, and the air film hole is inclined in the main flow direction by an angle of 10 ° to 90 °.
5. The gas turbine blade tip composite cooling structure combining sweeping impact and film hole as claimed in the preceding claims, wherein the ratio L/D of the length L of the gas turbine blade tip flat model to the hydraulic diameter D of the throat of the fluid vibration exciter is 1.5-200, the ratio W/D of the width W to the hydraulic diameter D of the throat of the fluid vibration exciter is 0.5-100, and the thickness H thereof is 03The ratio H of the hydraulic diameter D of the throat part of the fluid vibration exciter3the/D is 0.5-10.
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| Application Number | Priority Date | Filing Date | Title |
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| CN202111142033.6A CN113944516B (en) | 2021-09-28 | 2021-09-28 | Composite cooling structure for tip of gas turbine |
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| CN202111142033.6A CN113944516B (en) | 2021-09-28 | 2021-09-28 | Composite cooling structure for tip of gas turbine |
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| CN113944516B CN113944516B (en) | 2024-04-02 |
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Cited By (1)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN115020028A (en) * | 2022-07-15 | 2022-09-06 | 北京航空航天大学 | Superconducting cable with fluid oscillation structure for enhanced cooling |
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| EP1247940A1 (en) * | 1999-06-15 | 2002-10-09 | Mitsubishi Heavy Industries, Ltd. | Gas turbine stationary blade |
| CN103161513A (en) * | 2011-12-15 | 2013-06-19 | 通用电气公司 | Improved nozzle vane for a gas turbine engine |
| CN106014489A (en) * | 2016-07-15 | 2016-10-12 | 中国科学院工程热物理研究所 | Turbine blade provided with cooling structure, and gas turbine using turbine blade |
| CN107435563A (en) * | 2017-05-05 | 2017-12-05 | 西北工业大学 | A kind of case structure with tip clearance control and the flowing control of leaf top |
| CN108150224A (en) * | 2017-12-21 | 2018-06-12 | 西安交通大学 | A kind of eddy flow is the same as impacting cooling structure inside the turbine blade being combined |
| CN110700896A (en) * | 2019-11-29 | 2020-01-17 | 四川大学 | Gas turbine rotor blade with swirl impingement cooling structure |
| CN113153444A (en) * | 2021-04-09 | 2021-07-23 | 西安交通大学 | Turbine blade internal impingement cooling structure based on ultrasonic wave enhanced heat transfer |
| CN113236372A (en) * | 2021-06-07 | 2021-08-10 | 南京航空航天大学 | Gas turbine guide vane blade with jet oscillator and working method |
-
2021
- 2021-09-28 CN CN202111142033.6A patent/CN113944516B/en active Active
Patent Citations (8)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| EP1247940A1 (en) * | 1999-06-15 | 2002-10-09 | Mitsubishi Heavy Industries, Ltd. | Gas turbine stationary blade |
| CN103161513A (en) * | 2011-12-15 | 2013-06-19 | 通用电气公司 | Improved nozzle vane for a gas turbine engine |
| CN106014489A (en) * | 2016-07-15 | 2016-10-12 | 中国科学院工程热物理研究所 | Turbine blade provided with cooling structure, and gas turbine using turbine blade |
| CN107435563A (en) * | 2017-05-05 | 2017-12-05 | 西北工业大学 | A kind of case structure with tip clearance control and the flowing control of leaf top |
| CN108150224A (en) * | 2017-12-21 | 2018-06-12 | 西安交通大学 | A kind of eddy flow is the same as impacting cooling structure inside the turbine blade being combined |
| CN110700896A (en) * | 2019-11-29 | 2020-01-17 | 四川大学 | Gas turbine rotor blade with swirl impingement cooling structure |
| CN113153444A (en) * | 2021-04-09 | 2021-07-23 | 西安交通大学 | Turbine blade internal impingement cooling structure based on ultrasonic wave enhanced heat transfer |
| CN113236372A (en) * | 2021-06-07 | 2021-08-10 | 南京航空航天大学 | Gas turbine guide vane blade with jet oscillator and working method |
Cited By (1)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN115020028A (en) * | 2022-07-15 | 2022-09-06 | 北京航空航天大学 | Superconducting cable with fluid oscillation structure for enhanced cooling |
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| CN113944516B (en) | 2024-04-02 |
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