CN112302727A - Turbine blade leading edge cooling structure - Google Patents
Turbine blade leading edge cooling structure Download PDFInfo
- Publication number
- CN112302727A CN112302727A CN202011323799.XA CN202011323799A CN112302727A CN 112302727 A CN112302727 A CN 112302727A CN 202011323799 A CN202011323799 A CN 202011323799A CN 112302727 A CN112302727 A CN 112302727A
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- China
- Prior art keywords
- leading edge
- blade
- cooling structure
- turbine blade
- edge cooling
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- 238000001816 cooling Methods 0.000 title claims abstract description 46
- 230000009977 dual effect Effects 0.000 claims description 3
- 239000007789 gas Substances 0.000 description 6
- 230000000694 effects Effects 0.000 description 5
- 239000000112 cooling gas Substances 0.000 description 3
- 238000005192 partition Methods 0.000 description 2
- 230000009286 beneficial effect Effects 0.000 description 1
- 239000002737 fuel gas Substances 0.000 description 1
- 230000004927 fusion 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
- 239000007787 solid Substances 0.000 description 1
- 230000002194 synthesizing effect Effects 0.000 description 1
Images
Classifications
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
- F01D5/12—Blades
- F01D5/14—Form or construction
- F01D5/18—Hollow blades, i.e. blades with cooling or heating channels or cavities; Heating, heat-insulating or cooling means on blades
- F01D5/187—Convection cooling
-
- 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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- Engineering & Computer Science (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Turbine Rotor Nozzle Sealing (AREA)
Abstract
The invention discloses a cooling structure for the front edge of a turbine blade, wherein a cooling channel (2) and a double-swirl cavity (4) are arranged at the root part of the blade close to the front edge of the blade, the cooling channel (2) and the double-swirl cavity (4) are communicated through a plurality of swirl nozzles (3), a plurality of second air film holes (6) communicated with the double-swirl cavity (4) are arranged on a pressure surface (7) of the blade, and a plurality of first air film holes (5) communicated with the double-swirl cavity (4) are arranged on a suction surface (8) of the blade. The invention can improve the heat exchange nonuniformity of the front edge of the blade and has the characteristics of good heat exchange performance and low pressure loss.
Description
Technical Field
The invention relates to a cooling blade of a gas turbine, in particular to a cooling structure for a front edge of a turbine blade.
Background
With the increasing requirements of modern gas turbines on efficiency and output power, the inlet temperature of the turbine is increased, which can greatly increase the thermal load of turbine parts, and the development of novel efficient cooling structures is concerned with the performance of the gas turbine and the whole gas turbine.
The rotational flow cooling is a novel blade leading edge cooling mode, in the cooling mode, cold air enters a cooling chamber through a rotational flow nozzle in a tangential direction, large-scale high-speed rotary motion is formed, a target surface boundary layer is washed, and cold air mixing is enhanced to realize cooling of the blade. The rotational flow cooling mode has the characteristics of good cooling performance and low pressure loss. For impingement cooling, the heat transfer distribution is more uniform.
Disclosure of Invention
The invention aims to provide a turbine blade leading edge cooling structure for improving the cooling performance of the leading edge of a turbine blade.
In order to achieve the purpose, the invention is realized by the following technical scheme:
a cooling channel and a double-swirl cavity are arranged at the root part of a blade and are close to the front edge of the blade, the cooling channel and the double-swirl cavity are communicated through a plurality of swirl nozzles, a plurality of second air film holes communicated with the double-swirl cavity are arranged on the pressure surface of the blade, and a plurality of first air film holes communicated with the double-swirl cavity are arranged on the suction surface of the blade.
The invention is further improved in that the double vortex cavity is formed by overlapping two circular cavities.
The invention is further improved in that the two circular diameters of the chamber cross section of the double vortex cavity are equal.
The invention is further improved in that the diameter ratio of the circle center distance to the circle of the two circular chambers is more than 0.4.
The invention is further improved in that the cross section of the swirl nozzle is in a round corner rectangle shape.
The invention is further improved in that the incident position of the swirl nozzle is positioned at the midpoint of a connecting line of two circle centers of the double swirl cavities.
A further improvement of the invention is that the swirl nozzle is positioned at a location close to the tangential of the circular chamber of the pressure face of the vane.
A further improvement of the invention is that the incidence position of the swozzle is close to the tangential direction of the circular chamber of the suction surface of the blade.
The invention has the further improvement that the included angle between the axis of the air film hole of the pressure surface and the suction surface of the blade and the wall surface of the blade is 20-60 degrees, and the ratio of the distance of the air film hole to the diameter of the air film hole is 2.5-10.
The invention has at least the following beneficial technical effects:
according to the turbine blade leading edge cooling structure provided by the invention, the large-scale vortex flow is formed in the cooling chamber through the arranged vortex nozzles, the vortices of the two circular chambers can also mutually cross, the surface heat exchange is further enhanced, and the surface heat exchange is more uniform compared with other cooling modes.
To traditional whirl cooling structure, cooling gas gets into the whirl cavity after inlet channel or inlet port, follow blade leading edge passageway downward or upward flow behind the whirl in the cavity, the whirl hole department that is close to the exhaust end will bear strong transverse flow and hinder, the whirl cooling effect weakens relatively, traditional whirl cooling belongs to the category of forced convection heat transfer promptly, the joining in of first air film hole of whirl cavity and second air film hole, to the whirl cavity, the effect of having added some wall boundary layer suctions in other words, show to improve the whirl cooling effect of a plurality of whirl nozzles near the exhaust end to the wall, the air film cooling effect to the outer wall of blade has also been added simultaneously.
Furthermore, the double-vortex cavity in the front edge structure provided by the invention is in a double-circular fusion type, and the front edge structure is simple in structure and convenient to process.
Drawings
FIG. 1 is a partial cutaway view of a turbine blade leading edge cooling structure;
FIG. 2 is a view of a solid model sectioned from a cooling channel; wherein FIG. 2(b) is a sectional view taken along line A-A of FIG. 2 (a);
FIG. 3 is a schematic view showing a configuration of a swirler for a turbine blade leading edge cooling structure according to the present invention, and FIG. 3(a) shows a nozzle incident position at a midpoint of a connecting line between two centers of a circle of a dual swirl chamber; FIG. 3(b) shows the nozzle incident position located tangentially to the circular chamber near the pressure face; fig. 3(c) shows the nozzle incident position at the tangential direction of the circular chamber near the suction surface.
Detailed Description
The invention is described in further detail below with reference to the accompanying drawings:
as shown in fig. 1 to 3, in the cooling structure for the front edge of the turbine blade provided by the present invention, a cooling channel 2 and a double-swirl chamber 4 are disposed at the root of the blade near the front edge of the blade, the cooling channel 2 and the double-swirl chamber 4 are communicated through a plurality of swirl nozzles 3, a plurality of second film holes 6 communicated with the double-swirl chamber 4 are disposed on a pressure surface 7 of the blade, and a plurality of first film holes 5 communicated with the double-swirl chamber 4 are disposed on a suction surface 8 of the blade.
After entering a cooling channel 2 which is close to the front edge of the blade and is separated by a second partition plate 9 from the root of the blade, cooling gas is sprayed into a double-cyclone cavity 4 through a plurality of cyclone nozzles 3 arranged on a first partition plate 10, and after convective heat exchange is carried out in the double-cyclone cavity 4, the cooling gas enters main flow fuel gas through a second gas film hole 6 arranged on the pressure surface of the blade and a first gas film hole 5 arranged on the suction surface of the blade.
The double-vortex cavity 4 is formed by synthesizing two circular cavities, the cross section of the two circular cavities is a union set formed by intersecting two circles, the two circular diameters of the cross sections of the two circular cavities are equal, and the diameter ratio of the distance between the two circle centers and the circular diameter is larger than 0.4.
The section of the swirl nozzle 3 is a rounded rectangle, and the incident position is positioned at the midpoint of a connecting line of two circle centers of the double swirl cavities, and is close to the tangential direction of a circular cavity of the pressure surface 7 of the blade or the tangential direction of the circular cavity of the suction surface 8 of the blade.
And air film holes are respectively arranged on the blade pressure surface sides of the two circular rotational flow chambers close to the blade pressure surface 7 and the blade suction surface sides of the circular rotational flow chambers close to the blade suction surface, the included angles between the axes of the air film holes of the blade pressure surface and the blade suction surface and the wall surface of the blade are 20-60 degrees, and the ratio of the distance between the air film holes and the diameter of the air film holes is 2.5-10.
The length-width ratio of the rounded rectangular structure of the swirl nozzle 3 can be flexibly changed, and the number of the air film holes and the number of the swirl nozzles on the pressure surface and the suction surface of the blade can be freely changed according to the length of the blade.
The specific embodiments of the present invention are merely exemplary and should not be construed as limiting the scope of the invention, which is intended to cover all equivalent variations and modifications of the invention.
Claims (9)
1. The utility model provides a turbine blade leading edge cooling structure, its characterized in that, has seted up cooling channel (2) and two whirl chambeies (4) near the blade leading edge at the blade root, communicates through a plurality of whirl nozzles (3) between cooling channel (2) and two whirl chambeies (4), sets up a plurality of second air film holes (6) that communicate with two whirl chambeies (4) at blade pressure face (7), sets up a plurality of first air film holes (5) that communicate with two whirl chambeies (4) at blade suction face (8).
2. The turbine blade leading edge cooling structure as claimed in claim 1, wherein the dual vortex cavity (4) is formed by stacking two circular chambers.
3. A turbine blade leading edge cooling structure as claimed in claim 2, wherein the two circular diameters of the chamber cross-section of the dual vortex cavity (4) are equal.
4. The turbine blade leading edge cooling structure as claimed in claim 3, wherein the ratio of the distance between the centers of the two circular chambers and the diameter of the circle is greater than 0.4.
5. A turbine blade leading edge cooling structure as claimed in claim 2, wherein the cross-section of the swirler (3) is a rounded rectangle.
6. The turbine blade leading edge cooling structure as claimed in claim 5, wherein the incidence position of the swirler (3) is located at the midpoint of a line connecting two centers of circles of the double swirl chambers (4).
7. A turbine blade leading edge cooling structure as claimed in claim 5 wherein the incidence of the swozzle (3) is located tangentially to the circular chamber of the blade pressure face (7).
8. A turbine blade leading edge cooling structure as claimed in claim 5 wherein the incidence of the swozzle (3) is located tangentially to the circular chamber of the blade suction surface (8).
9. The turbine blade leading edge cooling structure as claimed in claim 1, wherein an included angle between an axis of a film hole of the pressure surface and the suction surface of the blade and a wall surface of the blade is 20-60 degrees, and a diameter ratio of a space of the film hole to the film hole is 2.5-10 degrees.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202011323799.XA CN112302727A (en) | 2020-11-23 | 2020-11-23 | Turbine blade leading edge cooling structure |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202011323799.XA CN112302727A (en) | 2020-11-23 | 2020-11-23 | Turbine blade leading edge cooling structure |
Publications (1)
| Publication Number | Publication Date |
|---|---|
| CN112302727A true CN112302727A (en) | 2021-02-02 |
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Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN202011323799.XA Pending CN112302727A (en) | 2020-11-23 | 2020-11-23 | Turbine blade leading edge cooling structure |
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| CN (1) | CN112302727A (en) |
Cited By (3)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN113006879A (en) * | 2021-03-19 | 2021-06-22 | 西北工业大学 | Aeroengine turbine film cooling hole with vortex generator |
| CN114109518A (en) * | 2021-11-29 | 2022-03-01 | 西安交通大学 | Turbine blade leading edge ribbed rotational flow-air film composite cooling structure |
| CN114215607A (en) * | 2021-11-29 | 2022-03-22 | 西安交通大学 | Turbine blade leading edge rotational flow cooling structure |
Citations (10)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US6099251A (en) * | 1998-07-06 | 2000-08-08 | United Technologies Corporation | Coolable airfoil for a gas turbine engine |
| GB0226291D0 (en) * | 2002-11-12 | 2002-12-18 | Rolls Royce Plc | Turbine components |
| US20060002795A1 (en) * | 2004-07-02 | 2006-01-05 | Siemens Westinghouse Power Corporation | Impingement cooling system for a turbine blade |
| US20060056967A1 (en) * | 2004-09-10 | 2006-03-16 | Siemens Westinghouse Power Corporation | Vortex cooling system for a turbine blade |
| US20160115796A1 (en) * | 2013-05-20 | 2016-04-28 | Kawasaki Jukogyo Kabushiki Kaisha | Turbine blade cooling structure |
| CN106761951A (en) * | 2017-01-23 | 2017-05-31 | 中国航发沈阳发动机研究所 | The leading edge cooling structure and the engine with it of a kind of turbine rotor blade |
| 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 |
| CN108979734A (en) * | 2018-07-18 | 2018-12-11 | 上海交通大学 | A kind of turbo blade multichannel cooling structure and device with eddy flow |
| CN111120008A (en) * | 2019-12-10 | 2020-05-08 | 西安交通大学 | Novel turbine blade rotational flow cooling structure |
| CN214366218U (en) * | 2020-11-23 | 2021-10-08 | 西安热工研究院有限公司 | Turbine blade leading edge cooling structure |
-
2020
- 2020-11-23 CN CN202011323799.XA patent/CN112302727A/en active Pending
Patent Citations (10)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US6099251A (en) * | 1998-07-06 | 2000-08-08 | United Technologies Corporation | Coolable airfoil for a gas turbine engine |
| GB0226291D0 (en) * | 2002-11-12 | 2002-12-18 | Rolls Royce Plc | Turbine components |
| US20060002795A1 (en) * | 2004-07-02 | 2006-01-05 | Siemens Westinghouse Power Corporation | Impingement cooling system for a turbine blade |
| US20060056967A1 (en) * | 2004-09-10 | 2006-03-16 | Siemens Westinghouse Power Corporation | Vortex cooling system for a turbine blade |
| US20160115796A1 (en) * | 2013-05-20 | 2016-04-28 | Kawasaki Jukogyo Kabushiki Kaisha | Turbine blade cooling structure |
| CN106761951A (en) * | 2017-01-23 | 2017-05-31 | 中国航发沈阳发动机研究所 | The leading edge cooling structure and the engine with it of a kind of turbine rotor blade |
| 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 |
| CN108979734A (en) * | 2018-07-18 | 2018-12-11 | 上海交通大学 | A kind of turbo blade multichannel cooling structure and device with eddy flow |
| CN111120008A (en) * | 2019-12-10 | 2020-05-08 | 西安交通大学 | Novel turbine blade rotational flow cooling structure |
| CN214366218U (en) * | 2020-11-23 | 2021-10-08 | 西安热工研究院有限公司 | Turbine blade leading edge cooling structure |
Non-Patent Citations (2)
| Title |
|---|
| 范小军;杜长河;周源远;李亮;丰镇平;: "复合冲击和复合旋流冷却特性的对比研究", 工程热物理学报, no. 12, 15 December 2018 (2018-12-15) * |
| 范小军;杜长河;李亮;李森;: "气膜孔几何位置对旋流冷却流动与传热特性的影响", 西安交通大学学报, no. 07, 21 April 2016 (2016-04-21) * |
Cited By (3)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN113006879A (en) * | 2021-03-19 | 2021-06-22 | 西北工业大学 | Aeroengine turbine film cooling hole with vortex generator |
| CN114109518A (en) * | 2021-11-29 | 2022-03-01 | 西安交通大学 | Turbine blade leading edge ribbed rotational flow-air film composite cooling structure |
| CN114215607A (en) * | 2021-11-29 | 2022-03-22 | 西安交通大学 | Turbine blade leading edge rotational flow cooling structure |
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