CN102953762A - Turbine airfoil and method for controlling a temperature of a turbine airfoil - Google Patents
Turbine airfoil and method for controlling a temperature of a turbine airfoil Download PDFInfo
- Publication number
- CN102953762A CN102953762A CN2012102819649A CN201210281964A CN102953762A CN 102953762 A CN102953762 A CN 102953762A CN 2012102819649 A CN2012102819649 A CN 2012102819649A CN 201210281964 A CN201210281964 A CN 201210281964A CN 102953762 A CN102953762 A CN 102953762A
- Authority
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- China
- Prior art keywords
- conduit
- plane
- airfoil
- pressure fluid
- contiguous
- 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.)
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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
- F01D11/00—Preventing or minimising internal leakage of working-fluid, e.g. between stages
- F01D11/02—Preventing or minimising internal leakage of working-fluid, e.g. between stages by non-contact sealings, e.g. of labyrinth type
- F01D11/04—Preventing or minimising internal leakage of working-fluid, e.g. between stages by non-contact sealings, e.g. of labyrinth type using sealing fluid, e.g. steam
-
- 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
- F01D11/00—Preventing or minimising internal leakage of working-fluid, e.g. between stages
- F01D11/005—Sealing means between non relatively rotating elements
- F01D11/006—Sealing the gap between rotor blades or blades and rotor
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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
- F05D2240/00—Components
- F05D2240/80—Platforms for stationary or moving blades
- F05D2240/81—Cooled platforms
-
- 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
- F05D2260/205—Cooling fluid recirculation, i.e. after cooling one or more components is the cooling fluid recovered and used elsewhere for other purposes
-
- 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
- F05D2260/221—Improvement of heat transfer
- F05D2260/2212—Improvement of heat transfer by creating turbulence
-
- 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
- F05D2260/221—Improvement of heat transfer
- F05D2260/2214—Improvement of heat transfer by increasing the heat transfer surface
- F05D2260/22141—Improvement of heat transfer by increasing the heat transfer surface using fins or ribs
Landscapes
- Engineering & Computer Science (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Turbine Rotor Nozzle Sealing (AREA)
Abstract
According to one aspect of the invention, a turbine airfoil (200) includes a platform (202) and a blade (204) extending from the platform (202). The airfoil (200) also includes a slot (208) formed in a slashface (210) of the platform (202), the slot (208) being configured to receive a pressurized fluid via passages (212) and configured to direct the pressurized fluid to a selected region of the turbine airfoil (200). The pressurized fluid forms a barrier to restrict the flow of hot gas accross the slashface (210), thus improving the airfoil life. And corresponding method for controlling the temperature of the airfoil.
Description
Technical field
Theme disclosed herein relates to turbine.More specifically, this theme relates to the airfoil (airfoil) in the turbine to be arranged in.
Background technique
In gas turbine engine, burner changes into heat energy with the chemical energy of fuel or air-fuel mixture.This heat energy is sent to the turbine that heat energy is changed into mechanical energy by the fluid from the air of compressor normally.Several factor affecting heat energy are to the efficient of mechanical transformation of energy.This factor can comprise blade passing frequency, fuel supply fluctuation, fuel type and reactivity, positive (head-on) volume of burner, fuel nozzle design, air-fuel profile, flame profile, air-fuel mixing, flame holding, combustion temperature, turbine component design, the dilution of hot gas path temperature and delivery temperature.For example, in the combustion temperatures such as the select location in burner and turbine nozzle district, can realize that the combustion efficiency and the power that improve generate.Sometimes, can shorten the life-span and increase wearing and tearing and the destruction of some member at the high-temperature of some burner and turbine zone.Therefore, expectation is the temperature of control in the turbine to reduce wear and to increase life-span of turbine component.
Summary of the invention
According to an aspect of the present invention, turbine airfoil comprises platform and the blade that extends from platform.This airfoil also comprises the conduit in the inclined-plane (slashface) that is formed at platform, this conduit be configured to via passage receive pressure fluid and be configured to direct pressurized fluid to the selection area of turbine airfoil to improve the airfoil life-span.
According to a further aspect in the invention, be provided for the method for cooled turbine airfoil spare, wherein, the method comprises the passage in the platform that makes the pressure fluid inflow be formed at turbine airfoil.The method also comprises makes pressure fluid flow into conduit the inclined-plane be formed at platform from passage, this conduit be configured to direct pressurized fluid to the selection area of turbine airfoil to improve the airfoil life-span.
These and other advantage and feature will become more apparent from following description by reference to the accompanying drawings.
Description of drawings
Be regarded as that theme of the present invention is pointed out especially and clearly advocate in the last claim of this specification.Of the present invention aforesaid apparent from following detailed description by reference to the accompanying drawings with further feature and advantage, wherein:
Fig. 1 is the embodiment's of gas turbine engine schematic diagram, and this gas turbine engine comprises burner, fuel nozzle, compressor and turbine;
Fig. 2 is the embodiment's of airfoil side view;
Fig. 3 is embodiment's the end elevation of the assembly of airfoil;
Fig. 4 is another embodiment's of airfoil perspective view;
Fig. 5 is the embodiment's of airfoil detailed end elevation; With
Fig. 6 is another embodiment's of airfoil detailed end elevation;
The mode of this detailed description by example with reference to the description of drawings embodiments of the invention together with advantage and feature.
List of parts
100 turbine systems
102 compressors
104 burners
106 turbines
108 axles
110 nozzles
112 fuel supplies
200 airfoils
202 platforms
204 blades
206 bottoms
208 conduits
210 inclined-planes
212 passages
214 hot gas paths
216 leading edges
219 trailing edges
220 cavitys
222 are used for the recess of pin
300 airfoils
302 platforms
304 blades
306 bottoms
308 conduits
310 inclined-planes
312 cavitys
314 hot gass stream
316 conduits
318 inclined-planes
320 conduits
322 inclined-planes
324 cooling fluids that receive
400 airfoils
402 platforms
404 blades
406 bottoms
408 conduits
410 inclined-planes
412 passages
414 notches
416 surfaces
418 streams
500 platforms
502 inclined-planes
504 conduits
600 platforms
602 inclined-planes
604 conduits.
Embodiment
Fig. 1 is the embodiment's of combustion gas turbine systems 100 schematic diagram.This system 100 comprises compressor 102, burner 104, turbine 106, axle 108 and fuel nozzle 110.In an embodiment, this system 100 can comprise a plurality of compressors 102, a plurality of burner 104, a plurality of turbine 106, a plurality of axle 108 and a plurality of fuel nozzle 110.As described, this compressor 102 and turbine 106 connect by axle 108.Axle 108 is single axle or be linked together to form a plurality of shaft parts of axle 108.
On the one hand, burner 104 utilizes liquid and/or gaseous fuel, such as the synthetic gas of rock gas or rich hydrogen, with operating turbine engines.For example, fuel nozzle 110 is communicated with forced air and fuel accommodating fluid from compressor 102.Fuel nozzle 110 is made air-fuel mixture, and discharges this air-fuel mixture to burner 104, thereby causes burning, and this burning produces the pressure exhaust of heat.The pressure exhaust of burner 104 guiding heat enters turbine nozzle (or " first order jet nozzle ") by transition piece, thereby leaves nozzle or stator blade and guide turbine rotor blade into or cause turbine 106 to rotate during blade when gas.The rotation of turbine 106 causes axle 108 rotation, thus when air inflow compressor 102 pressurized air.In an embodiment, airfoil (also being nozzle or movable vane) is arranged in the different piece (such as at compressor 102 or turbine 106) of turbine, because not identical temperature, the hot gas conductance of crossing airfoil in the different piece of this turbine causes wearing and tearing and the thermal fatigue of turbine components.The temperature of the part of control turbine airfoil can reduce wear and make in burner and can reach higher combustion temperature, thereby improves performance.The zone of the part of control such as airfoil and near the temperature in the zone the part to be improving component's life, discusses in detail following with reference to accompanying drawing 2 to 6.Although following discussion mainly concentrates on gas turbine, the concept of discussing is not limited to gas turbine.
Fig. 2 is the side view of the part of exemplary airfoil 200.This airfoil 200 comprises platform 202 and the blade 204 that extends from platform 202.Bottom 206 extension and can be used for fixedly airfoil to the part of rotor or stator, such as turbine wheel below platform 202.Conduit 208 is formed in the inclined-plane 210 of platform 202.Inclined-plane 210 is surfaces of platform, and it is configured to be adjacent to similar surface or the inclined-plane placement of contiguous airfoil.Several passages 212 are arranged in conduit and are configured to and will be communicated to conduit 208 such as the cooling fluid of pressurization or the fluid of the temperature control fluid of pressurization.The embodiment on inclined-plane 210 can comprise the single passage 212 of communication of fluid.In an embodiment, inclined-plane 210 is connected to contiguous inclined-plane and pressure fluid flows into conduit 208 to form fluid barriers, and it is configured to limit fluid and flows through the inclined-plane.In addition, along the mobile distribution type cooling that platform inclined-plane 202 is provided of the pressure fluid of conduit 208, thereby reduce wear and thermal fatigue and also improve simultaneously and prolong the airfoil life-span.
As described, hot gas path 214 flows to the trailing edge 218 of blade 204 from leading edge 216.Be formed at pressure fluid barriers restriction hot gass in the conduit 208 and cross flowing of the cavity 220 (be also referred to as " shank cavity ") of inclined-plane 210 in the bottom 206.The recess 222 that holds pin is positioned at below the platform 202.In an embodiment, pressure fluid also is configured to cool off recess 222 and territory, area in which the goods are in great demand.Cross inclined-plane 210 by restriction hot gas stream, the cooling fluid in the conduit 208 reduces wearing and tearing and the destruction on the bottom 206.In an embodiment, pressure fluid is the forced air for the selected part of cooling airfoil 200, and wherein, passage is used for the guiding cooling fluid to selected part.In addition, this passage can comprise passage 212, and wherein, pressure fluid was equipped with chill station 202 in 208 minutes by conduit.In this embodiment, conduit 208 comprises roughly semi-circular cross-section geometrical shape.As described, pressure fluid is configured to the direction of 214 streams along the hot gas path and flows, and wherein, fluid leaves the open trailing edge side of conduit 208.In other embodiments, the two end part of conduit 208 can be closed.Conduit 208 with end of closing can be configured to direct pressurized fluid to other zone of airfoil 200.In an embodiment, the high stress areas that the conduit 208 in inclined-plane 210 also can be such as the airfoil 200 of trailing edge 218 and platform 202 provides stress to eliminate, and wherein, conduit 208 weakens inclined-planes with from the high stress areas transfer load.As described, the cross-sectional geometry of conduit 208 is circle, an oval or avette part.In other embodiments, this cross-sectional geometry will comprise any suitable shape, such as triangle, rectangular or trapezoidal.In addition, conduit 208 can have the roughly the same cross section of crossing inclined-plane 210.Other embodiment can have the variable cross section for conduit 208, such as along the conduit 208 that changes aspect the sectional shape of its length or the size.For example, conduit 208 can have in one direction the sectional dimension that reduces forcing the fluting road 208 of wandering about as a refugee, or the size with increase is to reduce the flow velocity in the conduit outlet port.In another example, the shape transition that conduit 208 can be optimized from the heat transmission for 210 the part on the inclined-plane is to eliminating the shape that is optimized for the stress at another part place of 210 on the inclined-plane.
Aspect in, comprise that the turbine components of airfoil is formed by stainless steel or alloy, if wherein part the motor run duration suitably the cooling may be through heat fatigue.Should be noted in the discussion above that the apparatus and method for the temperature of controlling turbine components can be applicable to the cooling of the turbine rotor blade shown in Fig. 2 to 6, and the nozzle in turbogenerator, compressor stator blade or any other airfoil or hot gas path component.
Fig. 3 is the end elevation of the example components of airfoil 300 and airfoil 200.Airfoil 300 is roughly similar in appearance to airfoil 200 and comprise platform 302, blade 304 and bottom 306.Platform 302 is parts of airfoil body and comprises the conduit 308 that is formed in the inclined-plane 310.Inclined-plane 210 connects such as in the assembling of rotor or stator the time in turbine when airfoil 200,300 with being connected.Conduit 208 and 308 forms the cavity 312 that receives flow of pressurized fluid.Cavity 312 makes flow of pressurized fluid can control the temperature of platform 202 and 302.In addition, this cooling fluid barrier is formed in the cavity 312 and crosses inclined-plane 210 and 310 with restriction hot gas stream 314.In this embodiment, airfoil 200 and 300 comprises respectively the extra conduit 316 and 318 that is formed in inclined-plane 320 and 322.Inclined-plane 320 and 322 can be connected to the inclined-plane of contiguous airfoil.In an exemplary embodiment, passage 324 (also carrying work " path ") is arranged in the body of airfoil 200 and provides pressure fluid to flow to conduit 308 and passage 326 to conduit 208 and supply cooling fluid.Thereby the body of airfoil 200 can receive from the pressure fluid in source and via passage 324 and 326 supplied with pressurised fluid and arrive airfoil 300, thus the selection area of cooling airfoil 300.
Fig. 4 is the perspective view of the part of exemplary airfoil 400, and this airfoil 400 comprises platform 402, blade 404 and bottom 406.Platform 402 comprises the conduit 408 that is formed in the inclined-plane 410, is used for receiving the pressure fluid from passage 412.Platform 402 also comprises the feature such as notch 414, so that pressure fluid flows along the surface 416 of platform 402.Therefore, flow of pressurized fluid 418 is towards the open end of conduit 408 and pass notch 414.Pressure fluid in conduit 408 provides the distribution type cooling of platform 402 and forms barrier crosses inclined-plane 410 with limit fluid stream.By making pressure fluid flow through notch 414 also to the selection area such as surface 416, conduit 408 and notch 414 reduce thermal fatigue and wearing and tearing.Conduit 408 can comprise any suitable air-circulation features such as exemplary notch 414, and it utilizes structure, geometrical shape and/or passage to flow on such as the selected part of the airfoil of platform 402 with the guiding fluid and/or flows through selected part such as the airfoil of platform 402.Therefore, by fluid is guided on the surface 416 via notch 414, the temperature in surperficial 416 zones is controlled to reduce wear and thermal fatigue.In an embodiment, air-circulation features can comprise passage and/or notch, and it is configured to cooled region, such as blade 204,304,404 and/or bottom 206,306,406.
Fig. 5 and 6 is detailed end elevations of exemplary platform 500 and 600, and its utilization is respectively applied to conduit 502 and 602 different cross-sectional geometry.Exemplary geometrical shapes comprises semicircle, ellipse, trapezoidal and rectangular.Conduit 502 is included in the rectangular cross-sectional geometry in the inclined-plane 504, and wherein, this geometric configuration becomes to provide pressure fluid to flow to the selection area of platform 500.Similarly, conduit 602 is included in the trapezoid cross section geometrical shape in the inclined-plane 604.Therefore, conduit 208,308,408,502,602 the cross-sectional geometry selected part that the is constructed to airfoil fluid barriers that cooling is provided and/or forms selected volume flows with limit fluid.This conduit can be by forming such as any suitable method of casting and/or processing this platform.In addition, pressure fluid can be provided by outside and the dedicated source such as cooling fluid tank, the cooling-air that maybe can provide in inside for the other parts by turbine.Conduit and suitable cross-sectional geometry can be used for cooling off any turbine heat gas path component, and wherein, conduit provides cooling and/or limit fluid stream for member.In an embodiment, conduit be configured to direct pressurized fluid to the lower losses by mixture zone of airfoil to improve aerodynamic performance.For example, the bootable zone to airfoil of cooling fluid wherein, can not produce a large amount of turbulent flows when it runs into other flow such as hot gas.In an embodiment, cooling fluid is directed to the zone of airfoil so that energy can be from cooling fluid.This zone can comprise near the zone of throat of airfoil.
Although the present invention only describes in detail in conjunction with the embodiment of limited quantity, should be readily appreciated that, the invention is not restricted to this disclosed embodiment.On the contrary, the present invention can revise to comprise so far any amount of variation do not described, substitutes, displacement or the layout that is equal to, but they and the spirit and scope of the present invention match.Additionally, although various embodiment of the present invention describes, should be appreciated that aspect of the present invention can include only the embodiment of some descriptions.Therefore, the present invention is not considered and is limited to aforesaid description, but only limits to the scope of appended claims.
Claims (10)
1. a turbine airfoil (200), it comprises:
Platform (202);
Blade (204) from described platform (202) extension; With
Be formed at the conduit (208) in the inclined-plane (210) of described platform (202), described conduit (208) be configured to via passage (212) receive pressure fluid (418) and be configured to guide described pressure fluid (418) to the selection area of described turbine airfoil (200) to improve airfoil (200) life-span.
2. turbine airfoil according to claim 1 (200) is characterized in that, described conduit (208) be configured to guide described pressure fluid (418) to lower losses by mixture zone to improve aerodynamic performance.
3. turbine airfoil according to claim 1 (200), it is characterized in that, described blade (204) is configured to extend to hot gas path (214), and described conduit (208) is configured to utilize described pressure fluid (418) formation barrier to cross described inclined-plane (210) to the shank cavity with restriction hot gas stream.
4. turbine airfoil according to claim 1 (200), it is characterized in that, described conduit (208) is configured to be connected to the contiguous inclined-plane (310) of contiguous airfoil, and wherein, described contiguous inclined-plane (310) comprises contiguous conduit (308), to receive described pressure fluid (418) from the described passage (212) described inclined-plane (210).
5. turbine airfoil according to claim 1 (200), it is characterized in that, described conduit (208) is configured to be connected to the contiguous inclined-plane (310) of contiguous airfoil (300), and wherein, the described passage (212) in described inclined-plane (210) is configured to provide described pressure fluid (418) to described contiguous airfoil (300) via described contiguous inclined-plane (310).
6. turbine airfoil according to claim 5 (200) is characterized in that, described contiguous inclined-plane (210) comprises the have passage contiguous conduit (208) of (326), to receive described pressure fluid (418).
7. turbine airfoil according to claim 1 (200) is characterized in that, described conduit (208) comprises one the cross-sectional geometry that is selected from by in semicircle, the group that trapezoidal and rectangular forms.
8. method that is used for the temperature of control turbine airfoil (200), described method comprises:
Pressure fluid (418) is flowed into be formed at the passage (212) in the platform (202) of described turbine airfoil (200); And
Make pressure fluid (418) flow into conduit (208) the inclined-plane (210) be formed at described platform (202) from described passage (212), described conduit (208) be configured to guide described pressure fluid (418) to the selection area of described turbine airfoil (200) to improve airfoil (200) life-span.
9. method according to claim 8, it is characterized in that, make described pressure fluid (418) flow into described conduit (208) from described passage (212) and comprise, utilize described pressure fluid (418) to form barrier and cross described inclined-plane (210) with restriction hot gas stream.
10. method according to claim 8, it is characterized in that, comprise that the described conduit (208) that makes in described inclined-plane (210) is connected to the contiguous inclined-plane (310) of contiguous airfoil (300), wherein, described contiguous inclined-plane (310) comprises contiguous conduit (308), to receive described pressure fluid (418) from the described passage (212) described inclined-plane (210).
Applications Claiming Priority (2)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| US13/205,763 US20130039758A1 (en) | 2011-08-09 | 2011-08-09 | Turbine airfoil and method of controlling a temperature of a turbine airfoil |
| US13/205763 | 2011-08-09 |
Publications (1)
| Publication Number | Publication Date |
|---|---|
| CN102953762A true CN102953762A (en) | 2013-03-06 |
Family
ID=46639383
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN2012102819649A Pending CN102953762A (en) | 2011-08-09 | 2012-08-09 | Turbine airfoil and method for controlling a temperature of a turbine airfoil |
Country Status (3)
| Country | Link |
|---|---|
| US (1) | US20130039758A1 (en) |
| EP (1) | EP2557274A3 (en) |
| CN (1) | CN102953762A (en) |
Families Citing this family (1)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US10364680B2 (en) * | 2012-08-14 | 2019-07-30 | United Technologies Corporation | Gas turbine engine component having platform trench |
Citations (6)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US3816022A (en) * | 1972-09-01 | 1974-06-11 | Gen Electric | Power augmenter bucket tip construction for open-circuit liquid cooled turbines |
| US4940388A (en) * | 1988-12-07 | 1990-07-10 | Rolls-Royce Plc | Cooling of turbine blades |
| US5122033A (en) * | 1990-11-16 | 1992-06-16 | Paul Marius A | Turbine blade unit |
| US5531457A (en) * | 1994-12-07 | 1996-07-02 | Pratt & Whitney Canada, Inc. | Gas turbine engine feather seal arrangement |
| CN1782329A (en) * | 2004-11-24 | 2006-06-07 | 通用电气公司 | Controlled leakage pin and vibration damper |
| US20100316486A1 (en) * | 2009-06-15 | 2010-12-16 | Rolls-Royce Plc | Cooled component for a gas turbine engine |
Family Cites Families (8)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| JPH03213602A (en) * | 1990-01-08 | 1991-09-19 | General Electric Co <Ge> | Self cooling type joint connecting structure to connect contact segment of gas turbine engine |
| US6210111B1 (en) * | 1998-12-21 | 2001-04-03 | United Technologies Corporation | Turbine blade with platform cooling |
| US6955525B2 (en) * | 2003-08-08 | 2005-10-18 | Siemens Westinghouse Power Corporation | Cooling system for an outer wall of a turbine blade |
| US6984112B2 (en) * | 2003-10-31 | 2006-01-10 | General Electric Company | Methods and apparatus for cooling gas turbine rotor blades |
| US6887033B1 (en) * | 2003-11-10 | 2005-05-03 | General Electric Company | Cooling system for nozzle segment platform edges |
| US7524163B2 (en) * | 2003-12-12 | 2009-04-28 | Rolls-Royce Plc | Nozzle guide vanes |
| US8573942B2 (en) * | 2008-11-25 | 2013-11-05 | Alstom Technology Ltd. | Axial retention of a platform seal |
| GB0901129D0 (en) * | 2009-01-26 | 2009-03-11 | Rolls Royce Plc | Rotor blade |
-
2011
- 2011-08-09 US US13/205,763 patent/US20130039758A1/en not_active Abandoned
-
2012
- 2012-08-03 EP EP12179235.2A patent/EP2557274A3/en not_active Withdrawn
- 2012-08-09 CN CN2012102819649A patent/CN102953762A/en active Pending
Patent Citations (6)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US3816022A (en) * | 1972-09-01 | 1974-06-11 | Gen Electric | Power augmenter bucket tip construction for open-circuit liquid cooled turbines |
| US4940388A (en) * | 1988-12-07 | 1990-07-10 | Rolls-Royce Plc | Cooling of turbine blades |
| US5122033A (en) * | 1990-11-16 | 1992-06-16 | Paul Marius A | Turbine blade unit |
| US5531457A (en) * | 1994-12-07 | 1996-07-02 | Pratt & Whitney Canada, Inc. | Gas turbine engine feather seal arrangement |
| CN1782329A (en) * | 2004-11-24 | 2006-06-07 | 通用电气公司 | Controlled leakage pin and vibration damper |
| US20100316486A1 (en) * | 2009-06-15 | 2010-12-16 | Rolls-Royce Plc | Cooled component for a gas turbine engine |
Also Published As
| Publication number | Publication date |
|---|---|
| EP2557274A2 (en) | 2013-02-13 |
| EP2557274A3 (en) | 2017-05-17 |
| US20130039758A1 (en) | 2013-02-14 |
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Application publication date: 20130306 |