EP2620593A1 - Turbine airfoil and corresponding method of cooling - Google Patents
Turbine airfoil and corresponding method of cooling Download PDFInfo
- Publication number
- EP2620593A1 EP2620593A1 EP13152580.0A EP13152580A EP2620593A1 EP 2620593 A1 EP2620593 A1 EP 2620593A1 EP 13152580 A EP13152580 A EP 13152580A EP 2620593 A1 EP2620593 A1 EP 2620593A1
- Authority
- EP
- European Patent Office
- Prior art keywords
- film holes
- film
- turbine airfoil
- cooling
- helical
- 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.)
- Withdrawn
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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
- F01D5/187—Convection cooling
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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
- F05D2240/20—Rotors
- F05D2240/30—Characteristics of rotor blades, i.e. of any element transforming dynamic fluid energy to or from rotational energy and being attached to a rotor
- F05D2240/303—Characteristics of rotor blades, i.e. of any element transforming dynamic fluid energy to or from rotational energy and being attached to a rotor related to the leading edge of a rotor blade
-
- 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
- F05D2250/00—Geometry
- F05D2250/20—Three-dimensional
- F05D2250/25—Three-dimensional helical
-
- 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
- F05D2260/202—Heat transfer, e.g. cooling by film cooling
Definitions
- This application relates generally to gas turbine engines and, more particularly, to methods and apparatus for film cooling airfoils used within gas turbine engines.
- Gas turbine engines typically include a compressor, a combustor, and a turbine. Airflow entering the compressor is compressed and directed to the combustor where it is mixed with fuel and ignited, producing hot combustion gases used to drive the turbine. Blades and vanes used in the turbine section of a gas turbine engine each have an airfoil section that extends radially across an engine flowpath. During engine operation, the turbine blades and vanes are exposed to elevated temperatures that can lead to mechanical failure and corrosion. Therefore, it is common practice to make the blades and vanes from a temperature tolerant alloy and to apply corrosion resistant and thermally insulating coatings to the airfoil and other flowpath exposed surfaces. It is also widespread practice to cool the airfoils by flowing a coolant through the interior of the airfoils.
- a turbine vane or rotor blade typically includes a hollow airfoil, the outside of which is exposed to the hot combustion gases, and the inside of which is supplied with cooling fluid, which is typically compressed air.
- the airfoil includes leading and trailing edges, a pressure side, and a suction side.
- the pressure and suction sides connect at the airfoil leading and trailing edges, and span radially between an airfoil root and an airfoil tip.
- Film cooling holes extend between an internal cooling circuit defined within the airfoil and an outer surface of the airfoil. The film cooling holes route cooling fluid from the internal cooling circuit to the outside of the airfoil for film cooling the airfoil.
- Helical ribs in cooling holes have been used to generate a secondary flow pair of longitudinal vortices in the same direction as the turn of the rib. It may be desirable to utilize this known behavior in film holes to improve film cooling, increase coverage for film cooling and augment cooling efficiency.
- a turbine airfoil in a first aspect, includes a blade with a pressure sidewall and a suction sidewall joined together at chordally opposite leading and trailing edges, and at least one cooling hole disposed between the pressure sidewall and the suction sidewall adjacent the leading edge.
- a plurality of curved film holes extend between the at least one cooling hole and an exterior of the blade.
- a turbine airfoil in another aspect, includes a blade having a leading and trailing edges and an internal cooling circuit, and a plurality of film holes extending between the internal cooling circuit and an exterior of the blade.
- the plurality of film holes are shaped to generate a swirling flow exiting the film holes adjacent the leading edge to thereby enhance local convection and provide an insulating barrier to gaspath flow.
- the invention resides in a method of film cooling a turbine airfoil includes the steps of delivering cooling air to the internal cooling circuit, and flowing the cooling air from the internal cooling circuit through a plurality of film holes extending between the internal cooling circuit and an exterior of the blade.
- the flowing step comprises swirling the cooling air in the film holes and thereby providing an insulating barrier to gaspath flow.
- Film-cooling holes or film holes are widely used in modem gas turbines to cool the turbine airfoils that are exposed to hot combustion gases during operation of the turbine.
- the film-cooling holes provide cooling of the airfoil in several ways. Firstly, they provide film-cooling of the airfoil surface. Film-cooling is the cooling of a body or surface by maintaining a thin fluid layer over the affected area of a fluid that has a lower temperature than the operating environment. The fluid film insulates the film-cooled surface from the external operating environment, thereby reducing convective heat transfer from the external operating environment into the airfoil. Further, the film of the cooling fluid also removes heat from the airfoil surface.
- film-cooling also provides convective heat transfer from and cooling of the airfoil sidewall surrounding the film-cooling hole as the cooling air flows through it along the length of the hole.
- the film-cooling holes remove heat by providing an exhaust path for the cooling air that has been heated as it in turn cools the airfoil by passage through the airfoil cooling circuit.
- FIG. 1 shows a blade section 10 of a turbine airfoil.
- the blade includes a pressure sidewall 12 and a suction sidewall 14 joined together at chordally opposite leading 16 and trailing 18 edges.
- a cooling circuit is defined by a plurality of cooling passages or holes 20 that are disposed between the pressure sidewall 12 and the suction sidewall 14. At least one cooling hole 21 is positioned adjacent the leading edge 16.
- Film holes or film cooling holes are known that extend from one or more of the cooling holes 20 to an exterior of the blade.
- the film cooling holes are typically straight and direct cooling air from the cooling holes 20 to the blade exterior.
- the airfoil includes a plurality of curved film holes 22 that extend between the cooling hole 21 and the exterior of the blade 10. That is, a passage between the cooling hole 21 and the exterior of the blade 10 comprises a curved or twisted groove or the like such that air flowing through and exiting the film holes 22 is turning.
- An exemplary shape for the film holes may be helical, although other shapes may be contemplated, and the invention is not necessarily meant to be limited to the arrangement shown in the drawings.
- the film flow coming out of the holes no longer has a direct path, but rather exits in a swirling pattern, resulting in enhanced local convection with the holes as well as providing an insulating barrier to the gas path flow.
- the helical film holes 22 are oriented in both clockwise and counter-clockwise directions. Adjacent ones of the plurality of helical film holes 22 may thus be oriented in opposite directions. As a consequence of such structure, the exiting flow swirls in opposite vortices, further enhancing the advantageous effects of the design.
- a first group 24 of the helical film holes 22 may be oriented in one direction, while a second group 26 is oriented in an opposite direction. As shown, the first groups 24 and second groups 26 may alternate along a length of the blade 10.
- each of the first and second groups 24, 26 comprises three helical film holes 22.
- At least one of the helical film holes may comprise a double helical film hole 220. That is, the film hole 220 may comprise two (or more) interlaced helical grooves or passageways through which cooling air is passed.
- the cooling circuit with helical film holes serves to improve film cooling, increase coverage for film cooling and generally augment cooling efficiency.
- the swirling flow provides for enhanced local convection within the holes and also provides an insulating barrier to the gaspath flow.
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- Engineering & Computer Science (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Turbine Rotor Nozzle Sealing (AREA)
Abstract
Description
- This application relates generally to gas turbine engines and, more particularly, to methods and apparatus for film cooling airfoils used within gas turbine engines.
- Gas turbine engines typically include a compressor, a combustor, and a turbine. Airflow entering the compressor is compressed and directed to the combustor where it is mixed with fuel and ignited, producing hot combustion gases used to drive the turbine. Blades and vanes used in the turbine section of a gas turbine engine each have an airfoil section that extends radially across an engine flowpath. During engine operation, the turbine blades and vanes are exposed to elevated temperatures that can lead to mechanical failure and corrosion. Therefore, it is common practice to make the blades and vanes from a temperature tolerant alloy and to apply corrosion resistant and thermally insulating coatings to the airfoil and other flowpath exposed surfaces. It is also widespread practice to cool the airfoils by flowing a coolant through the interior of the airfoils.
- For example, a turbine vane or rotor blade typically includes a hollow airfoil, the outside of which is exposed to the hot combustion gases, and the inside of which is supplied with cooling fluid, which is typically compressed air. The airfoil includes leading and trailing edges, a pressure side, and a suction side. The pressure and suction sides connect at the airfoil leading and trailing edges, and span radially between an airfoil root and an airfoil tip. Film cooling holes extend between an internal cooling circuit defined within the airfoil and an outer surface of the airfoil. The film cooling holes route cooling fluid from the internal cooling circuit to the outside of the airfoil for film cooling the airfoil.
- Helical ribs in cooling holes have been used to generate a secondary flow pair of longitudinal vortices in the same direction as the turn of the rib. It may be desirable to utilize this known behavior in film holes to improve film cooling, increase coverage for film cooling and augment cooling efficiency.
- In a first aspect, the invention resides in a turbine airfoil includes a blade with a pressure sidewall and a suction sidewall joined together at chordally opposite leading and trailing edges, and at least one cooling hole disposed between the pressure sidewall and the suction sidewall adjacent the leading edge. A plurality of curved film holes extend between the at least one cooling hole and an exterior of the blade.
- In another aspect, the invention resides a turbine airfoil includes a blade having a leading and trailing edges and an internal cooling circuit, and a plurality of film holes extending between the internal cooling circuit and an exterior of the blade. The plurality of film holes are shaped to generate a swirling flow exiting the film holes adjacent the leading edge to thereby enhance local convection and provide an insulating barrier to gaspath flow.
- In yet another aspect, the invention resides in a method of film cooling a turbine airfoil includes the steps of delivering cooling air to the internal cooling circuit, and flowing the cooling air from the internal cooling circuit through a plurality of film holes extending between the internal cooling circuit and an exterior of the blade. The flowing step comprises swirling the cooling air in the film holes and thereby providing an insulating barrier to gaspath flow.
- Embodiments of the present invention will now be described, by way of example only, with reference to the accompanying drawings in which:
-
FIG. 1 is a perspective view of a blade portion of a turbine airfoil; -
FIG. 2 shows an exemplary arrangement of helical film holes; and -
FIG. 3 shows a double helical film hole. - Film-cooling holes or film holes are widely used in modem gas turbines to cool the turbine airfoils that are exposed to hot combustion gases during operation of the turbine. The film-cooling holes provide cooling of the airfoil in several ways. Firstly, they provide film-cooling of the airfoil surface. Film-cooling is the cooling of a body or surface by maintaining a thin fluid layer over the affected area of a fluid that has a lower temperature than the operating environment. The fluid film insulates the film-cooled surface from the external operating environment, thereby reducing convective heat transfer from the external operating environment into the airfoil. Further, the film of the cooling fluid also removes heat from the airfoil surface. Secondly, film-cooling also provides convective heat transfer from and cooling of the airfoil sidewall surrounding the film-cooling hole as the cooling air flows through it along the length of the hole. Thirdly, the film-cooling holes remove heat by providing an exhaust path for the cooling air that has been heated as it in turn cools the airfoil by passage through the airfoil cooling circuit.
-
FIG. 1 shows ablade section 10 of a turbine airfoil. The blade includes apressure sidewall 12 and asuction sidewall 14 joined together at chordally opposite leading 16 and trailing 18 edges. A cooling circuit is defined by a plurality of cooling passages orholes 20 that are disposed between thepressure sidewall 12 and thesuction sidewall 14. At least onecooling hole 21 is positioned adjacent the leadingedge 16. - Film holes or film cooling holes are known that extend from one or more of the
cooling holes 20 to an exterior of the blade. The film cooling holes are typically straight and direct cooling air from thecooling holes 20 to the blade exterior. With continued reference toFIG. 1 , and with reference toFIG. 2 , the airfoil according to preferred embodiments includes a plurality ofcurved film holes 22 that extend between thecooling hole 21 and the exterior of theblade 10. That is, a passage between thecooling hole 21 and the exterior of theblade 10 comprises a curved or twisted groove or the like such that air flowing through and exiting thefilm holes 22 is turning. An exemplary shape for the film holes may be helical, although other shapes may be contemplated, and the invention is not necessarily meant to be limited to the arrangement shown in the drawings. With the helical or other curved or twistedshaped film holes 22, the film flow coming out of the holes no longer has a direct path, but rather exits in a swirling pattern, resulting in enhanced local convection with the holes as well as providing an insulating barrier to the gas path flow. - Preferably, the
helical film holes 22 are oriented in both clockwise and counter-clockwise directions. Adjacent ones of the plurality ofhelical film holes 22 may thus be oriented in opposite directions. As a consequence of such structure, the exiting flow swirls in opposite vortices, further enhancing the advantageous effects of the design. As shown inFIG. 2 , in an exemplary embodiment, afirst group 24 of thehelical film holes 22 may be oriented in one direction, while a second group 26 is oriented in an opposite direction. As shown, thefirst groups 24 and second groups 26 may alternate along a length of theblade 10. In the embodiment shown inFIG. 2 , each of the first andsecond groups 24, 26 comprises threehelical film holes 22. - In yet another exemplary construction, with reference to
FIG. 3 , at least one of the helical film holes may comprise a doublehelical film hole 220. That is, thefilm hole 220 may comprise two (or more) interlaced helical grooves or passageways through which cooling air is passed. - The cooling circuit with helical film holes serves to improve film cooling, increase coverage for film cooling and generally augment cooling efficiency. The swirling flow provides for enhanced local convection within the holes and also provides an insulating barrier to the gaspath flow.
- While the invention has been described in connection with what is presently considered to be the most practical and preferred embodiments, it is to be understood that the invention is not to be limited to the disclosed embodiments, but on the contrary, is intended to cover various modifications and equivalent arrangements included within the spirit and scope of the appended claims.
Claims (11)
- A turbine airfoil comprising:a blade (10) including a pressure sidewall (12) and a suction sidewall (14) joined together at chordally opposite leading (16) and trailing (18) edges;at least one cooling hole (20) disposed between the pressure sidewall (12) and the suction sidewall (14) adjacent the leading edge (16); anda plurality of curved film holes (22) extending between the at least one cooling hole (20) and an exterior of the blade (10).
- A turbine airfoil according to claim 1, wherein the curved film holes (22) are helical.
- A turbine airfoil according to claim 2, wherein the plurality of helical film holes (22) are oriented in clockwise and counter-clockwise directions.
- A turbine airfoil according to claim 2 or 3, wherein adjacent ones of the plurality of helical film holes (22) are oriented in opposite directions.
- A turbine airfoil according to claim 2 or 3, wherein first groups (24) of the helical film holes (22) are oriented in one direction and second groups (26) of the helical film holes (22) are oriented in an opposite direction.
- A turbine airfoil according to claim 5, wherein the first groups (24) and second groups (28) alternate along a length of the blade (10).
- A turbine airfoil according to claim 5 or 6, wherein each of the first (24) and second (26) groups comprises three helical film holes (22).
- A turbine airfoil according to any of claims 2 to 7, wherein at least one of the plurality of helical film holes (22) comprises a double helical film hole (220).
- A turbine airfoil according to any preceding claim comprising:wherein the plurality of film holes (22) being shaped to generate a swirling flow exiting the film holes (22) adjacent the leading edge (16) to thereby enhance local convection and provide an insulating barrier to gaspath flow.
- A method of film cooling a turbine airfoil including a blade (10) with a leading edge (16) and a trailing edge (18) and having an internal cooling circuit (20), the method comprising:delivering cooling air to the internal cooling circuit (20); andflowing the cooling air from the internal cooling circuit (20) through a plurality of film holes (22) extending between the internal cooling circuit (20) and an exterior of the blade (10), the flowing step comprising swirling the cooling air in the film holes (22) and thereby providing an insulating barrier to gaspath flow.
- A method according to claim 10, wherein the plurality of film holes comprise helical film holes (22), and wherein the flowing step is practiced by flowing the cooling air from the internal cooling circuit (20) through the helical film holes (22).
Applications Claiming Priority (1)
Application Number | Priority Date | Filing Date | Title |
---|---|---|---|
US13/359,691 US20130195650A1 (en) | 2012-01-27 | 2012-01-27 | Gas Turbine Pattern Swirl Film Cooling |
Publications (1)
Publication Number | Publication Date |
---|---|
EP2620593A1 true EP2620593A1 (en) | 2013-07-31 |
Family
ID=47631309
Family Applications (1)
Application Number | Title | Priority Date | Filing Date |
---|---|---|---|
EP13152580.0A Withdrawn EP2620593A1 (en) | 2012-01-27 | 2013-01-24 | Turbine airfoil and corresponding method of cooling |
Country Status (5)
Country | Link |
---|---|
US (1) | US20130195650A1 (en) |
EP (1) | EP2620593A1 (en) |
JP (1) | JP2013155733A (en) |
CN (1) | CN103225517A (en) |
RU (1) | RU2013103432A (en) |
Families Citing this family (11)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
EP2971671B1 (en) * | 2013-03-15 | 2018-11-21 | United Technologies Corporation | Component, corresponding gas turbine engine and method of cooling a component |
EP2918782A1 (en) * | 2014-03-11 | 2015-09-16 | United Technologies Corporation | Component with cooling hole having helical groove and corresponding gas turbine engine |
US20170101870A1 (en) * | 2015-10-12 | 2017-04-13 | United Technologies Corporation | Cooling holes of turbine |
FR3052183B1 (en) * | 2016-06-02 | 2020-03-06 | Safran Aircraft Engines | TURBINE BLADE COMPRISING A COOLING AIR INTAKE PORTION INCLUDING A HELICOIDAL ELEMENT FOR SWIRLING THE COOLING AIR |
US20190003316A1 (en) * | 2017-06-29 | 2019-01-03 | United Technologies Corporation | Helical skin cooling passages for turbine airfoils |
US10787932B2 (en) * | 2018-07-13 | 2020-09-29 | Honeywell International Inc. | Turbine blade with dust tolerant cooling system |
JP7224928B2 (en) * | 2019-01-17 | 2023-02-20 | 三菱重工業株式会社 | Turbine rotor blades and gas turbines |
CN111140287B (en) * | 2020-01-06 | 2021-06-04 | 大连理工大学 | Laminate cooling structure adopting polygonal turbulence column |
CN111075510B (en) * | 2020-01-06 | 2021-08-20 | 大连理工大学 | Turbine blade honeycomb spiral cavity cooling structure |
CN112983561B (en) * | 2021-05-11 | 2021-08-03 | 中国航发四川燃气涡轮研究院 | Quincunx gas film hole and forming method, turbine blade and forming method and gas engine |
CN114876582B (en) * | 2022-06-28 | 2023-05-16 | 西北工业大学 | Turbine blade and aeroengine |
Citations (9)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
EP0641917A1 (en) * | 1993-09-08 | 1995-03-08 | United Technologies Corporation | Leading edge cooling of airfoils |
EP0971095A2 (en) * | 1998-07-06 | 2000-01-12 | United Technologies Corporation | A coolable airfoil for a gas turbine engine |
EP1013877A2 (en) * | 1998-12-21 | 2000-06-28 | United Technologies Corporation | Hollow airfoil for a gas turbine engine |
EP1759788A2 (en) * | 2005-09-01 | 2007-03-07 | United Technologies Corporation | Investment casting of cooled turbine airfoils |
EP1790821A1 (en) * | 2005-11-23 | 2007-05-30 | United Technologies Corporation | Refractory metal core cooling technologies for curved leading edge slots |
EP1847684A1 (en) * | 2006-04-21 | 2007-10-24 | Siemens Aktiengesellschaft | Turbine blade |
US20100119377A1 (en) * | 2008-11-12 | 2010-05-13 | Rolls-Royce Plc | Cooling arrangement |
US7785071B1 (en) * | 2007-05-31 | 2010-08-31 | Florida Turbine Technologies, Inc. | Turbine airfoil with spiral trailing edge cooling passages |
US7789626B1 (en) * | 2007-05-31 | 2010-09-07 | Florida Turbine Technologies, Inc. | Turbine blade with showerhead film cooling holes |
Family Cites Families (2)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US20090304494A1 (en) * | 2008-06-06 | 2009-12-10 | United Technologies Corporation | Counter-vortex paired film cooling hole design |
US8201621B2 (en) * | 2008-12-08 | 2012-06-19 | General Electric Company | Heat exchanging hollow passages with helicoidal grooves |
-
2012
- 2012-01-27 US US13/359,691 patent/US20130195650A1/en not_active Abandoned
-
2013
- 2013-01-24 EP EP13152580.0A patent/EP2620593A1/en not_active Withdrawn
- 2013-01-24 JP JP2013010707A patent/JP2013155733A/en active Pending
- 2013-01-25 RU RU2013103432/06A patent/RU2013103432A/en not_active Application Discontinuation
- 2013-01-25 CN CN201310029642XA patent/CN103225517A/en active Pending
Patent Citations (9)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
EP0641917A1 (en) * | 1993-09-08 | 1995-03-08 | United Technologies Corporation | Leading edge cooling of airfoils |
EP0971095A2 (en) * | 1998-07-06 | 2000-01-12 | United Technologies Corporation | A coolable airfoil for a gas turbine engine |
EP1013877A2 (en) * | 1998-12-21 | 2000-06-28 | United Technologies Corporation | Hollow airfoil for a gas turbine engine |
EP1759788A2 (en) * | 2005-09-01 | 2007-03-07 | United Technologies Corporation | Investment casting of cooled turbine airfoils |
EP1790821A1 (en) * | 2005-11-23 | 2007-05-30 | United Technologies Corporation | Refractory metal core cooling technologies for curved leading edge slots |
EP1847684A1 (en) * | 2006-04-21 | 2007-10-24 | Siemens Aktiengesellschaft | Turbine blade |
US7785071B1 (en) * | 2007-05-31 | 2010-08-31 | Florida Turbine Technologies, Inc. | Turbine airfoil with spiral trailing edge cooling passages |
US7789626B1 (en) * | 2007-05-31 | 2010-09-07 | Florida Turbine Technologies, Inc. | Turbine blade with showerhead film cooling holes |
US20100119377A1 (en) * | 2008-11-12 | 2010-05-13 | Rolls-Royce Plc | Cooling arrangement |
Also Published As
Publication number | Publication date |
---|---|
RU2013103432A (en) | 2014-07-27 |
US20130195650A1 (en) | 2013-08-01 |
CN103225517A (en) | 2013-07-31 |
JP2013155733A (en) | 2013-08-15 |
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