CN202023597U - Shrouded cooling blade of gas turbine - Google Patents
Shrouded cooling blade of gas turbine Download PDFInfo
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- CN202023597U CN202023597U CN2011201193785U CN201120119378U CN202023597U CN 202023597 U CN202023597 U CN 202023597U CN 2011201193785 U CN2011201193785 U CN 2011201193785U CN 201120119378 U CN201120119378 U CN 201120119378U CN 202023597 U CN202023597 U CN 202023597U
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- cooling
- integral shroud
- cooling chamber
- blade body
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Abstract
The utility model discloses a shrouded cooling blade of a gas turbine. A plurality of partition plates arrayed along radial direction are sequentially arranged on the inner wall of a blade body from the front edge of the blade body to the rear edge of the blade body, and an inner cavity of the blade body is divided into three return-flow cooling channels which are longitudinal along the blade body. Impingement cooling holes are arranged on the partition plates at the front end, and air film holes are arranged on the front edge of the blade body. The tail ends of the cooling channels in the middle of the blade body are connected with cooling cavities of a blade tip shroud, separating ribs, flow-disturbing columns and vent holes are arranged in the cooling cavities of the blade tip shroud. Venting gaps are arranged at the position of the rear edge of the blade body, and the flow-disturbing columns are located at the positions of the venting gaps. The shrouded cooling blade of the gas turbine has simple structure, improves cooling efficiency, reduces temperature of the blade body and the blade tip shroud while reducing air leakage loss of a blade tip and improves performances of the turbine blade and a complete machine.
Description
Technical field
The present invention relates to a kind of turbine cooling blade, be specially adapted to aeroengine and gas turbine.
Background technique
Turbine blade is one of critical component in the aviation gas turbine power plant, and its effect is that combustion gas heat energy is converted into the turbomachinery merit, and therefore, the quality of turbine blade design is directly connected to the performance of gas turbine and even whole motor.Because turbine blade of gas turbine is worked under high temperature, high pressure and high-revolving extreme environment, the thermal stress of environmental correclation, mechanical stress and ambient temperature will significantly reduce the working life of turbine blade therewith, how turbine blade is effectively cooled off, guarantee that normal, reliable, the long-term work of blade just seems particularly important.The turbine cooling blade technology can effectively reduce Turbine Blade Temperature Field, guarantee turbine blade and parts working life thereof, thereby the design of cooled blade has become one of core technology of gas turbine engine design.Be air loss that reduces turbine blade and the vibration characteristics of improving turbine blade, many turbine blade blade tips place has integral shroud.Consider that integral shroud and turbine blade joint will bear bigger centrifugal force during operation, turbine blade blade tip section area is less and near the fuel gas temperature of integral shroud is higher, the quality of integral shroud design will directly influence the service behaviour and the working life of turbine blade, thereby reduce the integral shroud temperature and quality is very important.The integral shroud cooling technology not only can reduce the temperature at integral shroud place, simultaneously integral shroud is emptied and reduced the integral shroud quality, alleviate the centrifugal force that is applied to integral shroud and turbine blade joint, when guaranteeing the blade working performance, prolonged the working life of turbine blade and parts thereof.At present, for example all do not relate to the integral shroud design in the turbine blade cooling structure form of proposition such as Chinese patent 200480001588.7,200710161323.9,200810056400.9,200820029544.0.Patent 200910003507.1 and 200910003508.6 has proposed turbine blade cooling integral shroud structural type, but this blade blade adopts the straight passage cooling, and blade position cold air utilization ratio is lower.Therefore, at shrouded turbines cooled blade characteristics, design efficient blade and integral shroud cooling structure form correspondingly is very important and urgent.
Summary of the invention
In order to overcome the not high deficiency of prior art shrouded turbines cooled blade unit cold air cooling effectiveness, the invention provides a kind of gas turbine band hat cooled blade, its structure is simple relatively, and under the condition that does not increase cooling air volume, can improve shrouded blade blade position cooling effectiveness, reduce the temperature at blade and integral shroud position effectively.
The technical solution adopted for the present invention to solve the technical problems is: comprise blade blade and integral shroud, blade blade inwall is aligned in sequence with the dividing plate that several are arranged along blade radial from blade inlet edge to trailing edge, its middle distance blade leading edge and trailing edge are not before two nearest dividing plates are divided into blade blade inner chamber, in, three independent sectors in back, all the other dividing plates with the blade inner chamber before, in, back three parts are divided into the cooling chamber of several connections respectively, vertically form three return flow type cooling channels along blade, the inlet of passage is all located blade root, cold air enters the blade cooling channel by blade root place cooling channel inlet, integral shroud is designed to hollow structure simultaneously, cold air can flow in the integral shroud cavity, forms the integral shroud cooling chamber.Impact cooling hole apart from having several on the nearest dividing plate of blade leading edge, blade leading edge inwall is formed impact cooling; The blade leading edge has several to penetrate into the air film hole of outer wall from the blade internal face simultaneously, after cold air flows out from air film hole, crooked downstream under main flow pressure and frictional force action, stick to the blade surface and form the air film cooling, blade wall and high-temperature fuel gas are separated, thereby reduce the blade temperature; The cooling chamber nearest apart from the blade trailing edge has exhaust seam at blade trailing edge place, penetrate into suction surface some turbulence columns that have been staggered at exhaust seam place from blade pressure surface, and turbulence columns has been strengthened the convection heat exchange at trailing edge position, improves cooling effectiveness; Blade top, cooling channel, middle part and integral shroud connecting place have some air outlet holes, and the cold air in the cooling channel, blade middle part of flowing through can directly enter the integral shroud cooling chamber by air outlet hole; Arrange the separation rib of some and perforation integral shroud cooling chamber substrate and top boards vertical in the integral shroud cooling chamber in blade central passage air outlet hole 1/3rd blade tip cross section chord length scopes with the blade tip string of a musical instrument, this separation rib does not link to each other with integral shroud cooling chamber sidewall, on separation rib, be furnished with simultaneously the some and rectangle fin separation rib Surface Vertical, this fin has increased separation rib surface heat transfer area, has improved cooling effectiveness; Arrange that outside blade central passage air outlet hole 1/3rd blade tip cross section chord length scopes plurality of rows runs through cooling chamber substrate and top board and staggered turbulence columns in the integral shroud cooling chamber, this turbulence columns has been strengthened the convection heat exchange in integral shroud cooling chamber downstream, further improves the integral shroud cooling effectiveness; Integral shroud cooling chamber sidewall is furnished with the plurality of rows pore apart from the extraneous position of chord length, a blade tip cross section of air outlet hole, blade middle part, and cold air is flowed through in cooling chamber and finally flowed out by exhaust port after separation rib and the turbulence columns.For the anterior cooling channel of blade, cooled gas finally flows out by blade leading edge air film hole along channel flow, forms the cooling on blade inlet edge surface; For cooling channel, blade middle part, cooled gas enters the integral shroud cooling chamber along channel flow by air outlet hole; For cooling channel, blade rear portion, cooled gas, finally flows out at trailing edge exhaust seam through the forced heat exchanging of the staggered turbulence columns in trailing edge place along channel flow; For the integral shroud cooling chamber, the flow through convection heat exchange of the staggered turbulence columns of separation rib and downstream that runs through cooling chamber substrate and top board of cooled gas finally flows out at integral shroud cooling chamber side exhaust port, forms complete path.
Described impact cooling hole is for circle and vertically run through dividing plate.
The invention has the beneficial effects as follows: 1., reduced the air loss of turbine blade blade tip, improved blade and overall efficiency owing to install integral shroud additional at turbine cooling blade blade tip place; 2. since the cooling channel be divided into before, during and after three independent sectors, have only cooling channel, blade middle part to link to each other with the integral shroud cooling chamber, make cooled gas all have the higher coefficient of heat transfer in blade front and rear edge position, improved cooling effect; 3. the blade cooling channel comprises a plurality of chambeies, and cooled gas has long stroke in blade, can increase blade position cooling effectiveness greatly; 4. be furnished with some rectangle fins on the separation rib in the integral shroud cooling chamber, and arrange turbulence columns, strengthened the convection heat exchange of cold air in the integral shroud cooling chamber, improve the integral shroud cooling effectiveness at the cold air downstream position; 5. the cooled blade structure is simple relatively, has good processability and future in engineering applications.
The present invention is further described below in conjunction with drawings and Examples.
Description of drawings
Fig. 1 is invention band hat cooled blade blade cross sectional representation;
Among the figure, 1-dividing plate one, 2-dividing plate two, 3-dividing plate three, 4-dividing plate four, 5-dividing plate five, 6-dividing plate six, 7-cooling chamber one, 8-cooling chamber two, 9-cooling chamber three, 10-cooling chamber four, 11-cooling chamber five, 12-cooling chamber six, 13-cooling chamber seven, 14-impacts cooling hole, 15-air film hole, 16-turbulence columns, 17-exhaust seam, A-blade outer wall, B-blade inwall.
Fig. 2 is invention band hat cooled blade blade cooling channel structural representation;
Among the figure, 18-inlet hole one, 19-inlet hole two, 20-inlet hole three, 21-inlet hole four, cooling channel, 22-blade middle part air outlet hole.
Fig. 3 is preced with the cooled blade monnolithic case for the invention band and cold air goes out flow diagram;
Fig. 4 is invention band hat cooled blade sectional drawing;
Fig. 5 is invention band hat cooling sole suction port design sketch;
Fig. 6 is invention band hat cooled blade integral shroud cooling chamber radial section figure;
Among the figure, 23-separation rib, 24-turbulence columns, 25-integral shroud cooling chamber exhaust port, 26-fin.
Embodiment
As depicted in figs. 1 and 2, band hat cooled blade blade comprises outer wall A and inwall B, is aligned in sequence with dividing plate 1 by the blade leading edge to trailing edge, dividing plate 22, dividing plate 33, dividing plate 44, dividing plate 55, dividing plate 66 is separated into cooling chamber 1 with the blade inner chamber, cooling chamber 28, cooling chamber 39, cooling chamber 4 10, cooling chamber 5 11, two independent sectors before, during and after cooling chamber 6 12 and cooling chamber 7 13, its central diaphragm 22 and dividing plate 55 are divided into cooled blade blade cooling channel; Equidistantly be arranged with 12 on the dividing plate 1 and impact cooling hole 14, when cooling blast flows at blade cooling chamber 28, by impact cooling hole ejection cooling air blade inwall B leading edge is blowed cooling air and cool off.The blade leading edge has staggered air film hole 15, penetrates into blade inwall A from blade inwall B, and as shown in Figure 3, a plurality of air film holes are arranged along the high direction of leaf in the blade leading edge, constitute the air film hole array.The air-flow that flows out from air film hole runs into the main flow high-temperature fuel gas, and turnover is flowed along blade outer wall A backward, and A has formed the film cooling in the blade outer wall.The cooling chamber 7 13 of blade trailing edge part forms some staggered turbulence columns 16 between blade pressure surface and suction surface, a plurality of turbulence columns are staggered along the high direction of leaf in last cooling chamber of blade, by the disturbance of turbulence columns to cooling blast in the cooling channel, influence flowing and conducting heat of air-flow, turbulence columns can also be strengthened the heat transfer of blade wall simultaneously.The upper end is interconnected with the upper end or lower end and lower end are interconnected between the adjacent blade cooling chamber, cold air can enter blade leading edge and cooling channel, blade middle part respectively from blade tenon bottom inflow hole 1 and inlet hole 2 19, can enter blade trailing edge cooling channel from inlet hole 3 20 and inlet hole 4 21, cold air flows out at leading edge air film hole 15 in the blade leading edge cooling channel, cold air flows into integral shroud cooling chamber shown in Figure 6 through cooling channel, blade middle part air outlet hole 22 in the cooling channel, blade middle part, and the cold air in the blade trailing edge cooling channel flows out at blade trailing edge exhaust seam 17.As shown in Figure 6, integral shroud is a hollow structure, and cold air can flow in cavity, forms the integral shroud cooling chamber.In blade central passage air outlet hole 1/22nd 3 blade tip cross section chord length scope, arrange that five arrange separation rib 23 vertical with the blade tip string of a musical instrument and perforation integral shroud cooling chamber substrate and top board in the integral shroud cooling chamber, this separation rib does not link to each other with integral shroud cooling chamber sidewall, on separation rib, be furnished with simultaneously the some and rectangle fin 26 separation rib Surface Vertical, arrange that outside blade central passage air outlet hole 1/3rd blade tip cross section chord length scopes plurality of rows runs through cooling chamber substrate and top board and staggered turbulence columns 24 in the integral shroud cooling chamber, integral shroud cooling chamber sidewall is furnished with four exhaust ports 25 apart from the extraneous position of chord length, a blade tip cross section of air outlet hole, blade middle part, and cold air flows into flow through behind the integral shroud cooling chamber separation rib and turbulence columns by blade central passage air outlet hole and finally flows out at integral shroud cooling chamber exhaust port.
As mentioned above, by band hat cooled blade design proposal of the present invention, can effectively reduce the air loss of turbine blade blade tip, under the prerequisite that does not increase cooling air delivery, improve simultaneously the cooling effect of key positions such as blade leading edge and trailing edge, improve the cooling effectiveness of blade cold air, cooling integral shroud design can alleviate the integral shroud quality simultaneously and reduce the integral shroud temperature, prolongs blade working life when guaranteeing turbine blade service behaviour and efficient, improves the gas turbine overall performance.
Claims (2)
1. gas combustion turbine cooling blade with hat, comprise blade blade and integral shroud, it is characterized in that: blade blade inwall is aligned in sequence with the dividing plate that several are arranged along blade radial from blade inlet edge to trailing edge, its middle distance blade leading edge and trailing edge are not before two nearest dividing plates are divided into blade blade inner chamber, in, three independent sectors in back, all the other dividing plates with the blade inner chamber before, in, back three parts are divided into the cooling chamber of several connections respectively, vertically form three return flow type cooling channels along blade, the inlet of passage is all located blade root, cold air enters the blade cooling channel by blade root place cooling channel inlet, integral shroud is designed to hollow structure simultaneously, cold air can flow in the integral shroud cavity, forms the integral shroud cooling chamber; Impact cooling hole apart from having several on the nearest dividing plate of blade leading edge, the blade leading edge has several to penetrate into the air film hole of outer wall from the blade internal face; The cooling chamber nearest apart from the blade trailing edge has exhaust seam at blade trailing edge place, penetrate into suction surface some turbulence columns that have been staggered at exhaust seam place from blade pressure surface; Blade top, cooling channel, middle part and integral shroud connecting place have some air outlet holes, and the cold air in the cooling channel, blade middle part of flowing through enters the integral shroud cooling chamber by air outlet hole; Arrange the separation rib of some and perforation integral shroud cooling chamber substrate and top boards vertical in the integral shroud cooling chamber in blade central passage air outlet hole 1/3rd blade tip cross section chord length scopes with the blade tip string of a musical instrument, this separation rib does not link to each other with integral shroud cooling chamber sidewall, is furnished with the some and rectangle fin separation rib Surface Vertical simultaneously on separation rib; Arrange that outside blade central passage air outlet hole 1/3rd blade tip cross section chord length scopes plurality of rows runs through cooling chamber substrate and top board and staggered turbulence columns in the integral shroud cooling chamber; Integral shroud cooling chamber sidewall is furnished with the plurality of rows pore apart from the extraneous position of chord length, a blade tip cross section of air outlet hole, blade middle part.
2. the gas combustion turbine cooling blade of band hat according to claim 1 is characterized in that: described impact cooling hole is for circle and vertically run through dividing plate.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN2011201193785U CN202023597U (en) | 2011-04-21 | 2011-04-21 | Shrouded cooling blade of gas turbine |
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| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN2011201193785U CN202023597U (en) | 2011-04-21 | 2011-04-21 | Shrouded cooling blade of gas turbine |
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| CN202023597U true CN202023597U (en) | 2011-11-02 |
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| Application Number | Title | Priority Date | Filing Date |
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| CN2011201193785U Expired - Fee Related CN202023597U (en) | 2011-04-21 | 2011-04-21 | Shrouded cooling blade of gas turbine |
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Cited By (8)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN102128055A (en) * | 2011-04-21 | 2011-07-20 | 西北工业大学 | Gas turbine cooling blade with crown |
| CN103291372A (en) * | 2012-03-01 | 2013-09-11 | 通用电气公司 | Turbine bucket with contoured internal rib |
| 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 |
| CN109441557A (en) * | 2018-12-27 | 2019-03-08 | 哈尔滨广瀚动力技术发展有限公司 | A kind of high-pressure turbine guide vane of the marine gas turbine with cooling structure |
| US10738700B2 (en) | 2016-11-16 | 2020-08-11 | General Electric Company | Turbine assembly |
| CN112112688A (en) * | 2019-06-21 | 2020-12-22 | 斗山重工业建设有限公司 | Turbine stator blade, turbine including the same, and gas turbine |
| CN114109516A (en) * | 2021-11-12 | 2022-03-01 | 中国航发沈阳发动机研究所 | Turbine blade end wall cooling structure |
| CN114370305A (en) * | 2022-01-25 | 2022-04-19 | 杭州汽轮动力集团有限公司 | Gas turbine stationary blade composite cooling structure |
-
2011
- 2011-04-21 CN CN2011201193785U patent/CN202023597U/en not_active Expired - Fee Related
Cited By (13)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN102128055A (en) * | 2011-04-21 | 2011-07-20 | 西北工业大学 | Gas turbine cooling blade with crown |
| CN103291372A (en) * | 2012-03-01 | 2013-09-11 | 通用电气公司 | Turbine bucket with contoured internal rib |
| US9127561B2 (en) | 2012-03-01 | 2015-09-08 | General Electric Company | Turbine bucket with contoured internal rib |
| CN103291372B (en) * | 2012-03-01 | 2016-12-28 | 通用电气公司 | Turbine blade with corrugated inner rib |
| US10738700B2 (en) | 2016-11-16 | 2020-08-11 | General Electric Company | Turbine assembly |
| 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 |
| CN109441557A (en) * | 2018-12-27 | 2019-03-08 | 哈尔滨广瀚动力技术发展有限公司 | A kind of high-pressure turbine guide vane of the marine gas turbine with cooling structure |
| CN109441557B (en) * | 2018-12-27 | 2024-06-11 | 哈尔滨广瀚动力技术发展有限公司 | High-pressure turbine guide vane of marine gas turbine with cooling structure |
| CN112112688A (en) * | 2019-06-21 | 2020-12-22 | 斗山重工业建设有限公司 | Turbine stator blade, turbine including the same, and gas turbine |
| CN112112688B (en) * | 2019-06-21 | 2023-02-17 | 斗山重工业建设有限公司 | Turbine stator blade, turbine including the same, and gas turbine |
| CN114109516A (en) * | 2021-11-12 | 2022-03-01 | 中国航发沈阳发动机研究所 | Turbine blade end wall cooling structure |
| CN114370305A (en) * | 2022-01-25 | 2022-04-19 | 杭州汽轮动力集团有限公司 | Gas turbine stationary blade composite cooling structure |
| CN114370305B (en) * | 2022-01-25 | 2024-02-20 | 杭州汽轮控股有限公司 | Composite cooling structure for turbine stator blades of gas turbine |
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| Date | Code | Title | Description |
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| C14 | Grant of patent or utility model | ||
| GR01 | Patent grant | ||
| C17 | Cessation of patent right | ||
| CF01 | Termination of patent right due to non-payment of annual fee |
Granted publication date: 20111102 Termination date: 20140421 |