CN110700895A - Gas turbine rotor blade with tip cooling structure - Google Patents
Gas turbine rotor blade with tip cooling structure Download PDFInfo
- Publication number
- CN110700895A CN110700895A CN201911197124.2A CN201911197124A CN110700895A CN 110700895 A CN110700895 A CN 110700895A CN 201911197124 A CN201911197124 A CN 201911197124A CN 110700895 A CN110700895 A CN 110700895A
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- CN
- China
- Prior art keywords
- blade
- groove
- air film
- turbine rotor
- gas turbine
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- Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
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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
- F01D5/186—Film cooling
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
- F01D5/12—Blades
- F01D5/14—Form or construction
- F01D5/18—Hollow blades, i.e. blades with cooling or heating channels or cavities; Heating, heat-insulating or cooling means on blades
- F01D5/187—Convection cooling
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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 rotor blade with a blade top cooling structure, which comprises a blade root, a blade profile with a blade leading edge and a blade trailing edge and a blade platform for connecting the blade root and the blade profile; the top of the blade profile is provided with a top sealing cover with a groove, the groove is internally provided with at least one air film hole and at least one bulge structure respectively, the bulge structures are distributed at intervals along the central connecting line direction of the air film hole, and gaps are reserved between the two sides of the bulge structures and the edges of the two sides of the groove; the coverage area of the air film is increased through the protruding structure, the flow direction of the air film is changed, the air film coverage effect of two sides of the groove surface of the blade top is enhanced, the groove surface of the blade top is cooled by the uniform and reasonable air film under the condition that the total cooling air amount is not increased, and the cooling effect of the top of the turbine blade is enhanced. And the flow separation in the groove caused by the convex structure can increase the flow mixing in the groove and reduce the leakage amount and the leakage loss of the blade top.
Description
Technical Field
The invention relates to the technical field of turbine rotor blades of gas turbines, in particular to a turbine rotor blade of a gas turbine with a blade top cooling structure.
Background
With the increasing temperature of the inlet gas of the turbine of the gas turbine, the thermal load environment of the high-temperature components of the turbine is worse. To ensure a reasonable life of the high temperature turbine blades, they need to be cooled efficiently. Among these, the cooling pattern is most complicated, particularly in the form of high temperature turbine rotor blades. Due to manufacturing characteristics, after the blade is subjected to precision casting of the internal cooling channel cavity, the top cover is welded on the top of the blade to complete the closing of the internal cooling circuit. Meanwhile, in the high-temperature gas flow channel of the turbine, because the gas leaks and flows at the blade top of the rotor blade, the blade top of the turbine rotor blade is easily oxidized at high temperature due to overhigh temperature or overlarge thermal stress, and the failure phenomena of cracks, even ablation and the like occur.
Disclosure of Invention
The invention aims to provide a gas turbine rotor blade with a blade top cooling structure, and aims to solve the problem that the blade top cooling effect of the existing gas turbine rotor blade is poor.
The technical scheme for solving the technical problems is as follows: a gas turbine rotor blade with a tip cooling structure includes a blade root, a blade airfoil having a leading edge and a trailing edge, and a blade platform connecting the blade root and the blade airfoil; a serpentine channel enhanced convection cooling circuit for cooling the top of the blade is arranged in the blade profile;
the blade profile top is provided with the notched top closing cap, be provided with at least one air film hole and at least one protruding structure in the recess respectively, just protruding structure is followed air film hole central connecting line direction interval distribution, protruding structure's both sides with all there is the clearance between the recess both sides edge.
Further, the connecting line direction of the central line of the convex structure is consistent with the direction of a curved arc line from the front edge of the blade to the tail edge of the blade.
Further, the ratio of the height of the protruding structures to the depth of the grooves is 0.1-0.3.
Further, the surface of the top cover is coated with an oxidation resistant coating and a heat insulating coating.
Further, the protruding structure is a rectangular structure, a hemispherical structure or a ribbed structure.
Further, the protruding structure and the top cover are of an integrated structure.
The invention has the following beneficial effects: according to the gas turbine rotor blade with the blade top cooling structure, the coverage area of an air film is enlarged through the protruding structure, the flow direction of the air film is changed, and the air film coverage effect on two sides of the groove surface of the blade top is enhanced. Under the condition of not increasing the total cooling air quantity, the blade top groove surface is cooled by a uniform and reasonable air film, and the cooling effect of the top of the turbine blade is enhanced. And the flow separation in the groove caused by the convex structure can increase the flow mixing in the groove and reduce the leakage amount and the leakage loss of the blade top.
Drawings
FIG. 1 is a schematic structural view of the present invention;
FIG. 2 is a schematic view of the protrusion structure and the air film hole of the present invention;
the reference numerals shown in fig. 1 to 2 are respectively expressed as: 1-blade root, 20-blade leading edge, 21-blade trailing edge, 3-blade platform, 4-groove, 5-top cover, 6-air film hole, 7-convex structure and 8-gap.
Detailed Description
The principles and features of this invention are described below in conjunction with the following drawings, the examples of which are set forth to illustrate the invention and are not intended to limit the scope of the invention.
As shown in fig. 1 to 2, a gas turbine rotor blade with tip cooling structure includes a blade root 1, a blade profile 2 having a blade leading edge 20 and a blade trailing edge 21, and a blade platform 3 connecting the blade root 1 and the blade profile 2; the blade profile 2 is internally provided with a serpentine channel enhanced convection cooling circuit for cooling the top of the blade.
The top of the blade profile 2 is provided with a top sealing cover 5 with a groove 4, the groove 4 is internally provided with at least one air film hole 6 and at least one protruding structure 7 respectively, the protruding structures 7 are distributed at intervals along the central connecting line direction of the air film hole 6, and gaps 8 are arranged between the two sides of the protruding structures 7 and the edges of the two sides of the groove 4. The structure of the convex structure 7 and the top sealing cover 5 are integrally processed; the tip cap 5 is manufactured separately and is mounted to the blade tip in a welded form, thereby forming an integral structure with the blade tip. The convex structures 7 are regular-shaped structures or irregular-shaped structures, and preferably, the convex structures 7 are rectangular structures, hemispherical structures or rib structures. The top cover 5 and the partition plate inside the blade form a transverse cooling channel in a serpentine channel enhanced convection cooling loop, and the transverse cooling channel is provided with at least one opening leading to a main flow of fuel gas outside the blade; the cooling gas in the transverse cooling channel flows from the blade leading edge 20 to the blade trailing edge 21, and finally flows into the main gas flow from the top film hole 6 and the opening of the transverse cooling channel. Through protruding structure 7 structure with the air film suppression in recess 4, the air film coverage of 4 face both sides (exhibition direction) of reinforcing leaf top recess makes 4 faces of leaf top recess obtain even reasonable air film cooling effect, has increased the heat transfer area in the recess 4 face, effectively reduces 4 local high temperatures of leaf top recess and local thermal stress, provides reasonable effectual thermal protection ability. And the flow separation in the groove 4 is caused by the structure of the convex structure 7, so that the flow mixing in the groove 4 is increased, and the leakage amount and the leakage loss of the blade top are reduced.
As the whole blade is bent from the front edge 20 of the blade to the tail edge 21 of the blade and presents a certain radian, in order to improve the uniform distribution performance of the cooling working medium in the surface of the blade top groove 4, in the invention, the connecting line direction of the central line of the bulge structure 7 is consistent with the bending direction of the front edge 20 of the blade to the tail edge 21 of the blade. The distribution direction of the convex structures 7 is the same as the curvature direction of the whole radian of the blade, so that the cooling medium in the groove 4 at the top of the blade is distributed along the curvature direction, the distribution uniformity of the cooling medium is improved, and the cooling effect on two sides of the groove 4 is better.
In order to better press the cooling air film in the groove 4 of the blade top, the ratio of the height of the convex structure 7 to the depth of the groove 4 is 0.1-0.3.
In order to further improve the thermal protection capability of the blade top groove 4 surface, in the invention, the surface of the top cover 5 is coated with an oxidation resistant coating and a thermal insulation coating. The coating is applied mainly on the gas side surface of the top cover 5.
The above description is only a preferred embodiment of the invention, and should not be taken as limiting the invention, and any modifications, equivalents, improvements and the like that are within the spirit and principle of the invention should be included in the scope of the invention.
Claims (6)
1. A gas turbine rotor blade with a tip cooling structure, comprising a blade root (1), a blade profile (2) with a leading edge (20) and a trailing edge (21) and a blade platform (3) connecting the blade root (1) and the blade profile (2); a snake-shaped channel enhanced convection cooling circuit for cooling the top of the blade is arranged in the blade profile (2);
blade profile (2) top is provided with top closing cap (5) that have recess (4), be provided with at least one air film hole (6) and at least one protruding structure (7) in recess (4) respectively, just protruding structure (7) are followed air film hole (6) central line direction interval distribution, the both sides of protruding structure (7) with all have clearance (8) between recess (4) both sides edge.
2. The gas turbine rotor blade with tip cooling structure according to claim 1, wherein the centerline direction of the convex structure (7) coincides with the curved camber line direction of the blade leading edge (20) to the blade trailing edge (21).
3. Gas turbine rotor blade with tip cooling according to claim 1, characterised in that the ratio between the height of the raised structure (7) and the depth of the groove (4) is 0.1-0.3.
4. The gas turbine rotor blade with tip cooling structure according to any one of claims 1 to 3, wherein the tip cover (5) surface is coated with an oxidation resistant coating and a thermal barrier coating.
5. Gas turbine rotor blade with tip cooling according to claim 4, characterised in that the raised structure (7) is a rectangular structure, a hemispherical structure or a ribbed structure.
6. The gas turbine rotor blade with tip cooling structure according to claim 5, wherein the projection structure (7) is a unitary structure with the tip cover (5).
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN201911197124.2A CN110700895A (en) | 2019-11-29 | 2019-11-29 | Gas turbine rotor blade with tip cooling structure |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN201911197124.2A CN110700895A (en) | 2019-11-29 | 2019-11-29 | Gas turbine rotor blade with tip cooling structure |
Publications (1)
| Publication Number | Publication Date |
|---|---|
| CN110700895A true CN110700895A (en) | 2020-01-17 |
Family
ID=69207997
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN201911197124.2A Pending CN110700895A (en) | 2019-11-29 | 2019-11-29 | Gas turbine rotor blade with tip cooling structure |
Country Status (1)
| Country | Link |
|---|---|
| CN (1) | CN110700895A (en) |
Cited By (1)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN112240229A (en) * | 2020-10-20 | 2021-01-19 | 西北工业大学 | A high-efficient cooling structure for turbine power blade top |
Citations (3)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN102102543A (en) * | 2011-03-11 | 2011-06-22 | 北京华清燃气轮机与煤气化联合循环工程技术有限公司 | Turbine rotor blade of gas turbine |
| CN102943694A (en) * | 2012-12-05 | 2013-02-27 | 沈阳航空航天大学 | Clapboard-type labyrinth structure for moving blade tip |
| US20170058678A1 (en) * | 2015-08-31 | 2017-03-02 | Siemens Energy, Inc. | Integrated circuit cooled turbine blade |
-
2019
- 2019-11-29 CN CN201911197124.2A patent/CN110700895A/en active Pending
Patent Citations (3)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN102102543A (en) * | 2011-03-11 | 2011-06-22 | 北京华清燃气轮机与煤气化联合循环工程技术有限公司 | Turbine rotor blade of gas turbine |
| CN102943694A (en) * | 2012-12-05 | 2013-02-27 | 沈阳航空航天大学 | Clapboard-type labyrinth structure for moving blade tip |
| US20170058678A1 (en) * | 2015-08-31 | 2017-03-02 | Siemens Energy, Inc. | Integrated circuit cooled turbine blade |
Cited By (1)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN112240229A (en) * | 2020-10-20 | 2021-01-19 | 西北工业大学 | A high-efficient cooling structure for turbine power blade top |
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| PB01 | Publication | ||
| PB01 | Publication | ||
| SE01 | Entry into force of request for substantive examination | ||
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| RJ01 | Rejection of invention patent application after publication | ||
| RJ01 | Rejection of invention patent application after publication |
Application publication date: 20200117 |