CN112610284A - Gas turbine blade with spiral band - Google Patents
Gas turbine blade with spiral band Download PDFInfo
- Publication number
- CN112610284A CN112610284A CN202011497824.6A CN202011497824A CN112610284A CN 112610284 A CN112610284 A CN 112610284A CN 202011497824 A CN202011497824 A CN 202011497824A CN 112610284 A CN112610284 A CN 112610284A
- Authority
- CN
- China
- Prior art keywords
- blade
- gas turbine
- cooling
- gas
- air conditioning
- Prior art date
- 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.)
- Pending
Links
- 239000007789 gas Substances 0.000 claims abstract description 33
- 239000000112 cooling gas Substances 0.000 claims abstract description 16
- 210000003041 ligament Anatomy 0.000 claims abstract description 7
- 238000001816 cooling Methods 0.000 abstract description 26
- 230000000694 effects Effects 0.000 abstract description 11
- 238000004378 air conditioning Methods 0.000 abstract description 6
- 238000005406 washing Methods 0.000 abstract description 3
- 230000003014 reinforcing effect Effects 0.000 abstract description 2
- 230000002708 enhancing effect Effects 0.000 description 3
- 230000009286 beneficial effect Effects 0.000 description 2
- 230000014759 maintenance of location Effects 0.000 description 2
- 238000012986 modification Methods 0.000 description 2
- 230000004048 modification Effects 0.000 description 2
- 230000008646 thermal stress Effects 0.000 description 2
- 230000004888 barrier function Effects 0.000 description 1
- 230000000903 blocking effect Effects 0.000 description 1
- 239000002131 composite material Substances 0.000 description 1
- 230000007774 longterm Effects 0.000 description 1
- 239000000463 material Substances 0.000 description 1
- 230000002035 prolonged effect Effects 0.000 description 1
- 238000009991 scouring Methods 0.000 description 1
- 230000035882 stress 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
-
- 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
Landscapes
- Engineering & Computer Science (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Turbine Rotor Nozzle Sealing (AREA)
Abstract
The invention relates to a turbine blade of a gas turbine with helical ligaments, comprising: the blade leading edge, the blade trailing edge, blade suction surface, the blade pressure face, the blade top, the blade root, the vortex rib, the vortex post, the gas film hole, cooling gas access channel, the whirl cavity, whirl nozzle and spiral link, further strengthen the rotary motion of air conditioning in the whirl cavity through the spiral link, strengthen the washing away of air conditioning to the wall, the reinforcing is to the disturbance of wall boundary layer, improve the heat transfer effect, the time that extension air conditioning stayed in cooling channel, improve the utilization ratio of cooling gas, thereby reduce required air conditioning volume and reduce gas turbine's pneumatic punishment, improve gas turbine's efficiency, the inhomogeneous technical problem of current gas turbine blade leading edge part cooling has been solved.
Description
Technical Field
The invention relates to the technical field of turbine blades of gas turbines, in particular to a turbine blade of a gas turbine with spiral bands.
Background
The inlet air temperature of the H-grade gas turbine can reach more than 1600 ℃, and the inlet air temperature far exceeds the limit which can be borne by blade materials. The first stages of turbine blades are directly exposed to high-temperature gas and are washed by the high-temperature high-speed gas to bear severe heat load, and particularly, the working conditions of the region where the surface of the front edge of the blade is located are worse. At present, the commonly adopted cooling scheme of the front edge of the gas turbine comprises air film cooling and impingement cooling, and the impingement cooling is seriously influenced by cross flow. Effective impingement cooling needs to be combined with film cooling to avoid the negative effects of cross flow on internal heat transfer and to provide additional protection for surfaces downstream of the film holes. However, the film cooling has the problems of local stress concentration increase, extra pneumatic punishment, blocking of film holes under the long-term work of an industrial environment and the like.
The rotational flow cooling is a new cooling mode for the front edge of the turbine blade of the gas turbine at the present stage, and the rotational flow cooling and the impact air film composite cooling have equivalent cooling effect and are more effective than the single impact cooling. As shown in fig. 1, a cooling gas inlet hole is formed in the blade root 6 and connected with a cooling gas channel 10, a group of swirl nozzles 12 are arranged on the cooling gas channel 10 in the direction corresponding to the swirl chamber 11, cooling gas enters the group of swirl nozzles 12 with a small area from the cooling gas inlet hole of the blade root 6, the flow velocity of the cooling gas is further increased, then the cooling gas tangentially enters the swirl chamber 11, and forms a rotating flow along with the guiding of the surface of the swirl chamber 11, the high-speed jet strongly washes the wall surface, the boundary layer is damaged, and a relatively excellent heat transfer effect is realized.
However, since the rotationally flowing cold air tends to flow more axially than the circumferential rotational movement, the rotational flow movement is sufficient only on the front edge surface corresponding to the nozzle, and the effect of the rotational flow washing is significantly weakened in the region between the two nozzles, resulting in uneven cooling effect and higher thermal stress.
Disclosure of Invention
Aiming at the technical problems in the prior art, the spiral band and the turbine blade of the gas turbine are creatively and organically combined, the turbine blade of the gas turbine with the spiral band is designed, the spiral band is used for further enhancing the rotary motion of cold air in the vortex chamber, enhancing the scouring of the cold air on the wall surface, enhancing the disturbance on the boundary layer of the wall surface, improving the heat exchange effect and solving the technical problem of uneven cooling of the front edge part of the turbine blade of the gas turbine in the prior art.
The technical scheme for realizing the invention is as follows: a gas turbine blade having a helical ligament comprising: blade leading edge 1, blade trailing edge 2, blade suction face 3, blade pressure face 4, blade top 5, blade root 6, vortex rib 7, vortex post 8, gas film hole 9, cooling gas admission passage 10, whirl chamber 11, swirl nozzle 12, characterized by: it still includes: the spiral band 13 is arranged in the cyclone chamber 11, and the upper end and the lower end of the spiral band 13 are fixedly connected with the top and the bottom of the cyclone chamber 11 respectively.
The pitch of the spiral link 13 is 2 mm-5 mm.
The invention discloses a gas turbine blade with spiral bands, which has the beneficial effects that:
1. a gas turbine blade with spiral ties can prolong the retention time of cold air in a cooling channel, improve the utilization rate of cooling gas, reduce the required cold air amount, reduce the pneumatic punishment of a gas turbine and improve the efficiency of the gas turbine by adding the spiral ties;
2. the utility model provides a gas turbine blade with spiral tie, further strengthens the rotary motion of air conditioning in the whirl cavity, strengthens the washing away of air conditioning to the wall, and the reinforcing is to the disturbance of wall boundary layer, improves the heat transfer effect, and the tie can act as the effect of fin, further improves the heat transfer efficiency of blade, has solved the cooling effect uneven, produces the condition of higher thermal stress.
Drawings
FIG. 1 is a three-dimensional schematic view of a prior art gas turbine blade;
FIG. 2 is a three-dimensional schematic view of a gas turbine blade having helical ligaments;
in the figure: 1. the blade comprises a blade front edge, 2 blade tail edges, 3 blade suction surfaces, 4 blade pressure surfaces, 5 blade tops, 6 blade roots, 7 turbulence ribs, 8 turbulence columns, 9 gas film holes, 10 cooling gas inlet channels, 11 swirl chambers, 12 swirl nozzles and 13 spiral ties.
Detailed Description
The present invention will be described in further detail with reference to fig. 2 and the detailed description, which is provided herein for the purpose of illustration only and is not intended to limit the present invention.
As shown in FIG. 2, a gas turbine blade having a helical ligament comprising: blade leading edge 1, blade trailing edge 2, blade suction face 3, blade pressure face 4, blade top 5, blade root 6, vortex rib 7, vortex post 8, gas film hole 9, cooling gas admission passage 10, whirl cavity 11, whirl nozzle 12 and spiral tie 13 whirl cavity 11 in set up spiral tie 13, spiral tie 13 upper and lower both ends link firmly with 11 tops of whirl cavity and bottom respectively, spiral tie 13 the pitch be 3 mm.
The cooling air for cooling the leading edge of the turbine blade of a gas turbine enters the cooling air inlet channel 10 from the cooling air inlet opening of the blade root 6. Between the cooling gas inlet channel 10 and the swirl chamber 11 there is a barrier, on one side of which a set of swirl nozzles 12 are arranged, which are distributed equidistantly. The outer edge surface of the swirl chamber 11 is the same as the outer surface shape of the leading edge 1 of the turbine blade of the gas turbine, and the distance between the outer edge surface of the swirl chamber 11 and the surface of the leading edge 1 of the blade is equal everywhere, so that the arrangement of the swirl chamber is prevented from being limited by the shape of the blade, and the uniformity of cooling is ensured. After entering the air inlet channel 10, the cold air enters the group of swirl nozzles 12 with smaller area, further increase of the cold air flow speed is realized, and then the cold air tangentially enters the swirl chamber 11, impacts the spiral band 13 in the swirl chamber 11, and forms a rotating flow in the swirl chamber 11 under the spiral guide of the spiral band 13. The spiral ties 13 further promote the rotation of the cooling air in the swirl chamber to ensure better cooling of the leading edge surface between the two swirl nozzles 12. And the turbulence intensity of the internal cold air is enhanced, and the disturbance to the boundary layer is stronger. When the rotary motion of the cold air is increased, the retention time of the cold air in the cyclone chamber 11 is prolonged to a certain extent, so that the cold air can be fully utilized, the using amount of the cold air can be reduced, and the efficiency of the gas turbine is further improved. Furthermore, the presence of the helical ligaments 13 may act as fins, which may be beneficial to some extent in improving the heat transfer effect, and finally the cooling gas is discharged from the film holes 13 at the blade tip.
The foregoing is only a preferred embodiment of the present invention, and it should be noted that, for those skilled in the art, various modifications and improvements can be made without departing from the principle of the present invention, and these modifications and improvements should also be considered as the protection scope of the present invention.
Claims (2)
1. A gas turbine blade having a helical ligament comprising: blade leading edge (1), blade trailing edge (2), blade suction face (3), blade pressure side (4), blade top (5), root of blade (6), vortex rib (7), vortex post (8), film hole (9), cooling gas access channel (10), whirl chamber (11), swirl nozzle (12), characterized by: it still includes: the spiral band (13) is arranged in the cyclone chamber (11), and the upper end and the lower end of the spiral band (13) are fixedly connected with the top and the bottom of the cyclone chamber (11) respectively.
2. A turbine blade for a gas turbine according to claim 1, characterised in that the pitch of the helical ligament (13) is between 2 mm and 5 mm.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202011497824.6A CN112610284A (en) | 2020-12-17 | 2020-12-17 | Gas turbine blade with spiral band |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202011497824.6A CN112610284A (en) | 2020-12-17 | 2020-12-17 | Gas turbine blade with spiral band |
Publications (1)
| Publication Number | Publication Date |
|---|---|
| CN112610284A true CN112610284A (en) | 2021-04-06 |
Family
ID=75240261
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN202011497824.6A Pending CN112610284A (en) | 2020-12-17 | 2020-12-17 | Gas turbine blade with spiral band |
Country Status (1)
| Country | Link |
|---|---|
| CN (1) | CN112610284A (en) |
Cited By (2)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN115247575A (en) * | 2022-05-12 | 2022-10-28 | 中国航发四川燃气涡轮研究院 | Spiral turbine blade cooling unit and cooling structure |
| WO2024097458A1 (en) * | 2022-11-03 | 2024-05-10 | Ge Infrastructure Technology Llc | Turbine nozzle or blade with impingement cooling structure having thermal flex elements |
Citations (12)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| DE3306894A1 (en) * | 1983-02-26 | 1984-08-30 | MTU Motoren- und Turbinen-Union München GmbH, 8000 München | Turbine stator or rotor blade with cooling channel |
| US5002460A (en) * | 1989-10-02 | 1991-03-26 | General Electric Company | Internally cooled airfoil blade |
| US5704763A (en) * | 1990-08-01 | 1998-01-06 | General Electric Company | Shear jet cooling passages for internally cooled machine elements |
| US5993156A (en) * | 1997-06-26 | 1999-11-30 | Societe Nationale D'etude Et De Construction De Moteurs D'aviation Snecma | Turbine vane cooling system |
| US20040096313A1 (en) * | 2002-11-12 | 2004-05-20 | Harvey Neil W. | Turbine components |
| US20140140860A1 (en) * | 2012-01-20 | 2014-05-22 | Rolls-Royce Plc | Aerofoil cooling |
| US20150159494A1 (en) * | 2013-12-10 | 2015-06-11 | Gilles Carrier | Multi-orifice plate for cooling flow control in vane cooling passage |
| US20160312624A1 (en) * | 2013-12-20 | 2016-10-27 | United Technologies Corporation | Gas turbine engine component cooling cavity with vortex promoting features |
| 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 |
| US20190292918A1 (en) * | 2016-06-02 | 2019-09-26 | Safran Aircraft Engines | Turbine vane including a cooling-air intake portion including a helical element for swirling the cooling air |
| US20190330987A1 (en) * | 2018-04-25 | 2019-10-31 | United Technologies Corporation | Spiral cavities for gas turbine engine components |
| CN110700896A (en) * | 2019-11-29 | 2020-01-17 | 四川大学 | Gas turbine rotor blade with swirl impingement cooling structure |
-
2020
- 2020-12-17 CN CN202011497824.6A patent/CN112610284A/en active Pending
Patent Citations (12)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| DE3306894A1 (en) * | 1983-02-26 | 1984-08-30 | MTU Motoren- und Turbinen-Union München GmbH, 8000 München | Turbine stator or rotor blade with cooling channel |
| US5002460A (en) * | 1989-10-02 | 1991-03-26 | General Electric Company | Internally cooled airfoil blade |
| US5704763A (en) * | 1990-08-01 | 1998-01-06 | General Electric Company | Shear jet cooling passages for internally cooled machine elements |
| US5993156A (en) * | 1997-06-26 | 1999-11-30 | Societe Nationale D'etude Et De Construction De Moteurs D'aviation Snecma | Turbine vane cooling system |
| US20040096313A1 (en) * | 2002-11-12 | 2004-05-20 | Harvey Neil W. | Turbine components |
| US20140140860A1 (en) * | 2012-01-20 | 2014-05-22 | Rolls-Royce Plc | Aerofoil cooling |
| US20150159494A1 (en) * | 2013-12-10 | 2015-06-11 | Gilles Carrier | Multi-orifice plate for cooling flow control in vane cooling passage |
| US20160312624A1 (en) * | 2013-12-20 | 2016-10-27 | United Technologies Corporation | Gas turbine engine component cooling cavity with vortex promoting features |
| US20190292918A1 (en) * | 2016-06-02 | 2019-09-26 | Safran Aircraft Engines | Turbine vane including a cooling-air intake portion including a helical element for swirling the cooling air |
| 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 |
| US20190330987A1 (en) * | 2018-04-25 | 2019-10-31 | United Technologies Corporation | Spiral cavities for gas turbine engine components |
| CN110700896A (en) * | 2019-11-29 | 2020-01-17 | 四川大学 | Gas turbine rotor blade with swirl impingement cooling structure |
Cited By (3)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN115247575A (en) * | 2022-05-12 | 2022-10-28 | 中国航发四川燃气涡轮研究院 | Spiral turbine blade cooling unit and cooling structure |
| CN115247575B (en) * | 2022-05-12 | 2024-05-03 | 中国航发四川燃气涡轮研究院 | Helical turbine blade cooling unit and cooling structure |
| WO2024097458A1 (en) * | 2022-11-03 | 2024-05-10 | Ge Infrastructure Technology Llc | Turbine nozzle or blade with impingement cooling structure having thermal flex elements |
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| Date | Code | Title | Description |
|---|---|---|---|
| PB01 | Publication | ||
| PB01 | Publication | ||
| SE01 | Entry into force of request for substantive examination | ||
| SE01 | Entry into force of request for substantive examination | ||
| RJ01 | Rejection of invention patent application after publication | ||
| RJ01 | Rejection of invention patent application after publication |
Application publication date: 20210406 |