EP2093381A1 - Aube rotorique ou statorique de turbine à plateforme refroidie - Google Patents

Aube rotorique ou statorique de turbine à plateforme refroidie Download PDF

Info

Publication number
EP2093381A1
EP2093381A1 EP08003397A EP08003397A EP2093381A1 EP 2093381 A1 EP2093381 A1 EP 2093381A1 EP 08003397 A EP08003397 A EP 08003397A EP 08003397 A EP08003397 A EP 08003397A EP 2093381 A1 EP2093381 A1 EP 2093381A1
Authority
EP
European Patent Office
Prior art keywords
platform
turbine component
trailing edge
component according
cooling medium
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Withdrawn
Application number
EP08003397A
Other languages
German (de)
English (en)
Inventor
Vitali Bregman
Vladimir Filippov
Sergey Shukin
Esa Utriainen
Current Assignee (The listed assignees may be inaccurate. Google has not performed a legal analysis and makes no representation or warranty as to the accuracy of the list.)
Siemens AG
Original Assignee
Siemens AG
Priority date (The priority date 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 date listed.)
Filing date
Publication date
Application filed by Siemens AG filed Critical Siemens AG
Priority to EP08003397A priority Critical patent/EP2093381A1/fr
Priority to PCT/EP2009/051914 priority patent/WO2009106464A1/fr
Publication of EP2093381A1 publication Critical patent/EP2093381A1/fr
Withdrawn legal-status Critical Current

Links

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
    • F01D9/00—Stators
    • F01D9/02—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles
    • F01D9/04—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles forming ring or sector
    • F01D9/041—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles forming ring or sector using 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
    • F01D11/00—Preventing or minimising internal leakage of working-fluid, e.g. between stages
    • F01D11/001—Preventing or minimising internal leakage of working-fluid, e.g. between stages for sealing space between stator blade and rotor
    • 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
    • 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
    • F05D2240/10—Stators
    • F05D2240/12—Fluid guiding means, e.g. vanes
    • F05D2240/122—Fluid guiding means, e.g. vanes related to the trailing edge of a stator vane
    • 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
    • F05D2240/304—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 trailing 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
    • 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
    • F05D2260/201—Heat transfer, e.g. cooling by impingement of a fluid
    • 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
    • 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

Definitions

  • the invention relates to a turbine component, in particular a vane or a blade, comprising a platform, an airfoil portion extending outwardly from said platform and a shank portion extending inwardly from said platform, whereby said platform comprising a end zone forming a platform trailing edge, whereby inwardly orientated surface of said trailing edge is being ventilated by a flow of a cooling medium streaming outwardly along said surface.
  • Aforesaid turbine component is known from EP 1 512 835 A2 .
  • the trailing edge of a turbine blade platform is very difficult to cool today. Increasing turbine inlet temperatures make it necessary to cool the blade platform more efficiently than it is done today. Blade platform side faces are cooled by coolant leakage flows today but the platform trailing edge should be cooled more intentionally.
  • a cooled rotor blade for a gas turbine is known, which has a cooled airfoil portion, a shaft portion and a platform with a trailing edge.
  • An analogous technical solution is known from a cooled rotor blade disclosed in DE 12 47 072 .
  • cooled blade airfoil portion is cooled only. The platform and the trailing edge are not cooled and therefore they may suffer overheating damage.
  • US 2007/0009359 A1 discloses a gas turbine blade with film cooling channels permitting air to flow from a shank cavity through the platform exterior to the airfoil.
  • mentioned film cooling channels can be used also to purge a trailing edge undercut as disclosed in US 2005/0135936 A1 .
  • the combination of these documents misses any teaching for design of precise channel duct.
  • EP 1 512 835 A2 discloses a gas turbine blade comprising a platform, an airfoil portion extending outwardly from said platform and a shank portion extending inwardly from said platform, whereby said platform comprising a rear side forming a platform trailing edge, and whereby inwardly orientated surface of said trailing edge is being ventilated by a flow of a cooling medium streaming outwardly along said surface.
  • the stream of cooling medium discharges from an exhaust hole situated on the rear side of the shank portion. Said exhaust hole is fed by a channel connecting exhaust hole with an internal cavity provided within the shaft portion and supplied with coolant.
  • the exhaust hole is directed towards the inwardly orientated surface of said trailing edge, to permit edge ventilating flow of cooling medium to stream from exhaust hole towards said surface.
  • edge ventilating flow of cooling medium to stream from exhaust hole towards said surface.
  • the problem is solved by associating said inwardly orientated surface with at least one surface enlargement element.
  • a basic idea of this invention is to enhance the heat transfer between platform surface and the coolant by promoting coolant flow turbulence and thus increase the heat transfer coefficient and heat flux.
  • Surface enlargement elements according to this invention act as turbulators to promote coolant flow turbulence. Therefore, an improved cooling is achieved. Due to the fact that platform trailing edge temperature is reduced, this results in an increase of the lifetime and reliability of the blade or vane.
  • Surface enlargement elements that could be used for this purpose include: a plurality of dimples depressed into surface or pimples extending from surface. Extending elements could either be pin-shaped fins or ribs with rectangular cross section, both spaced to another. A preferred embodiment is equipped with ribs designed in a criss-cross pattern. This would also give additional mechanical strength. All these heat transfer surface enlargement elements may also be combined to further enhance heat transfer.
  • enlargement element extending from surface is a hook-shaped portion of trailing edge directed to the axis of the turbine.
  • this embodiment is combined with channel fed exhaust holes, whereby the axis of said hole and/or channel is inclined to the trailing edge and directed to the end zone of the trailing edge.
  • jet impingement cooling of the end part of the trailing edge which has, as a rule, the most severe thermal conditions and the highest probability of overheating if no cooling).
  • Enlargement element in sense of this invention may also be a perforated plate mounted upstream from said surface regarding flow of cooling medium, said plate extending substantially parallel to said surface, whereby stream of cooling medium is flowing through perforation of plate.
  • the plate could be attached to the shaft portion of the component to be cooled or to the shaft portion of the adjacent component.
  • the perforation in the plate acts as additional impingement cooling for the bottom of the platform thus improving cooling of the trailing edge.
  • This perforated plate is expected to increase the thermal flux about 20 to 40 percent.
  • the guide vane 1, shown in Fig. 1 has the airfoil portion 2 and the platform 3 with the trailing edge 4.
  • Platform 3 is connected with the shaft portion 5.
  • a cavity 6 within the shaft portion 2 below the platform 3 is connected with coolant (as a rule, cooling air) source 7.
  • the shaft portion 5 is provided with channel-like exhaust holes 8 connected with the cavity 6.
  • the axis 9 of this holes respectively of the channel are directed to the end zone 10 of the trailing edge.
  • the end zone of the trailing edge 4 has a end portion shaped like a hook 11, directed to the axis A of the turbine. There is jet impingement cooling of the hook shaped end portion of the trailing edge 4.
  • the example chosen for illustration is a guide vane 1, the same can be applied to a rotor blade of a gas turbine.
  • the gas turbine works in such a way that hot gas flows through a channel between the guide vane 1, getting a circumferential component of its velocity. Then said velocity provides the power of a turbine, exerting the circumferential force of the blades.
  • the coolant flows from the source 7 into the inner cavity 6 cooling the platform 3. Then the coolant discharges through the exhaust hole 8 as a flow F towards the bottom surface of platform 3 to be cooled. Since the axis 9 of the hole 8 is directed to the end zone of the trailing edge 4, this provides jet impingement cooling of the trailing edge. Said cooling results in effective cooling of the edge and increases the lifetime and reliability of the vane and the turbine.
  • the hook-shaped end portion 11 in zone 10 of the inward orientated surface of the platform 3 promotes a better air jet braking, thus increases the jet impingement effectiveness.
  • FIG. 2 shows yet another embodiment of the present invention relating to a blade 12 of a gas or steam turbine.
  • the coolant is not fed from within the shank portion but rather from a rotor/stator cavity 13 defined between the shank portion 5 of the blade 12 and the shank portion 14 of the adjacent vane 15. Since the pressure in rotor/stator cavity 13 is higher than the pressure in the gas / steam path 16, the coolant flow F is coming up from the rotor/stator cavity 13 passing platform trailing edge 4.
  • the latter is equipped with a plurality of surface enlargements elements 17, extending from the inwardly orientated surface of the trailing edge 4. These elements 17 improve heat transfer between the blade surface and the coolant flow F coming up from the cavity 13.
  • the elements act also as turbulators further increasing effectiveness of impingement cooling.
  • Figures 3a to 3e show miscellaneous embodiments of surface enlargement elements 17, each in sectional view as well as in perspective view.
  • a plurality of dimples 18 are depressed into inwardly orientated surface of the platform 3. Contrary to dimples a plurality of pimples may extend from surface.
  • Extending elements could either be pin-shaped fins 19 ( Figure 3b ) or ribs 20 with rectangular cross section ( Figure 3c and 3d ), both spaced to another.
  • ribs 20 are directed transverse to the flow F of the coolant.
  • Figure 3d ribs 20 are arranged interrupted and directed longitudinal to the flow F of the coolant.
  • a preferred embodiment is equipped with ribs 20 designed in a criss-cross pattern as shown in Figure 3e . This would also give additional mechanical strength. All these heat transfer surface enlargement elements may also be combined to further enhance heat transfer.
  • FIG. 4 shows a third embodiment of the invention characterized in a perforated plate 21.
  • This plate 21 is mounted to the shank portion 5 of blade 12 in a manner, that it extends substantially parallel to bottom surface of blade platform 3 to be cooled.
  • plate 21 is bent in a way following the face of shank portion 5.
  • flow F of cooling medium said plate is arranged upstream. The stream of cooling medium is flowing through the perforation of the plate 21.
  • perforation in the plate acts as additional impingement cooling for the bottom of the platform.
  • This perforated plate is expected to increase the thermal flux about 20 to 40 percent.
  • Figure 5 shows a forth embodiment of the invention characterized in a perforated plate 21.
  • the shape of plate 21 shown in Figure 5 is completely plain. It is associated to the platform trailing edge 4 of blade 12 in manner, that it extends substantially parallel to bottom surface of blade platform 3 to be cooled. On its right side it is suitably fixed to the shank portion 14 of the vane 15, e. g. by welding. On its left side it contacts shank portion 5 of blade 12 by a sealing element 22 of some sort to minimize leakage of coolant between the stationary plate 21 and the rotating rotor surface. Eventually, perforated plate 21 mounted upstream to be flushed with coolant effects similar to dimples 18 shown in Figure 3a .

Landscapes

  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
EP08003397A 2008-02-25 2008-02-25 Aube rotorique ou statorique de turbine à plateforme refroidie Withdrawn EP2093381A1 (fr)

Priority Applications (2)

Application Number Priority Date Filing Date Title
EP08003397A EP2093381A1 (fr) 2008-02-25 2008-02-25 Aube rotorique ou statorique de turbine à plateforme refroidie
PCT/EP2009/051914 WO2009106464A1 (fr) 2008-02-25 2009-02-18 Aube ou ailette de turbine avec plate-forme de refroidissement

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
EP08003397A EP2093381A1 (fr) 2008-02-25 2008-02-25 Aube rotorique ou statorique de turbine à plateforme refroidie

Publications (1)

Publication Number Publication Date
EP2093381A1 true EP2093381A1 (fr) 2009-08-26

Family

ID=39537462

Family Applications (1)

Application Number Title Priority Date Filing Date
EP08003397A Withdrawn EP2093381A1 (fr) 2008-02-25 2008-02-25 Aube rotorique ou statorique de turbine à plateforme refroidie

Country Status (2)

Country Link
EP (1) EP2093381A1 (fr)
WO (1) WO2009106464A1 (fr)

Cited By (13)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
WO2011029420A1 (fr) * 2009-09-10 2011-03-17 Mtu Aero Engines Gmbh Dispositif de déflexion pour un écoulement de fuite dans une turbine à gaz et turbine à gaz correspondante
CN102953772A (zh) * 2011-08-08 2013-03-06 通用电气公司 用于控制涡轮机中的流的***和方法
DE102013200497A1 (de) * 2013-01-15 2014-07-17 Siemens Aktiengesellschaft Drosseleinheit sowie Rotor und Gasturbine
FR3001492A1 (fr) * 2013-01-25 2014-08-01 Snecma Stator de turbomachine avec controle passif de la purge
FR3028883A1 (fr) * 2014-11-25 2016-05-27 Snecma Arbre de rotor de turbomachine comportant une surface d'echange thermique perfectionnee
EP3101236A1 (fr) * 2015-06-01 2016-12-07 United Technologies Corporation Joints de plate-forme de bord de fuite
EP3273004A1 (fr) * 2016-07-22 2018-01-24 General Electric Company Refroidissement d'aube de turbine
EP3521570A1 (fr) * 2018-02-05 2019-08-07 United Technologies Corporation Composant et moteur à turbine à gaz associé
EP3521571A1 (fr) * 2018-01-31 2019-08-07 United Technologies Corporation Refroidissement par impact de plate-forme des aubes de turbine
US10619484B2 (en) 2015-01-22 2020-04-14 General Electric Company Turbine bucket cooling
US10822987B1 (en) 2019-04-16 2020-11-03 Pratt & Whitney Canada Corp. Turbine stator outer shroud cooling fins
FR3107919A1 (fr) * 2020-03-03 2021-09-10 Safran Aircraft Engines Aube creuse de turbomachine et plateforme inter-aubes équipées de saillies perturbatrices de flux de refroidissement
FR3107920A1 (fr) * 2020-03-03 2021-09-10 Safran Aircraft Engines Aube creuse de turbomachine et plateforme inter-aubes équipées de saillies perturbatrices de flux de refroidissement

Families Citing this family (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US9068513B2 (en) * 2013-01-23 2015-06-30 Siemens Aktiengesellschaft Seal assembly including grooves in an inner shroud in a gas turbine engine
WO2015112227A2 (fr) * 2013-11-12 2015-07-30 United Technologies Corporation Multiples trous d'injection pour ailette de moteur à turbine à gaz

Citations (18)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
DE1247072B (de) 1962-12-05 1967-08-10 Gen Motors Corp Hohlschaufel, insbesondere fuer Gasturbinen
DE2639511A1 (de) 1975-09-08 1977-03-17 Gen Electric Kuehlluftleckstromausnutzung
FR2417639A1 (fr) * 1976-05-14 1979-09-14 Rolls Royce Dispositif refrigerant pour aube de distributeur de moteur a turbine a gaz
DE3248161A1 (de) 1981-12-28 1983-07-07 United Technologies Corp., 06101 Hartford, Conn. Kuehlbare schaufel
GB2244520A (en) * 1990-05-31 1991-12-04 Gen Electric Nozzle assembly for a gas turbine engine
US5201847A (en) * 1991-11-21 1993-04-13 Westinghouse Electric Corp. Shroud design
US5954475A (en) * 1996-01-08 1999-09-21 Mitsubishi Jukogyo Kabushiki Kaisha Gas turbine stationary blade
EP1380726A2 (fr) * 2002-07-10 2004-01-14 Mitsubishi Heavy Industries, Ltd. Aube statorique pour turbine à gaz et turbine à gaz comprenant cet élément
WO2004038179A1 (fr) * 2002-10-24 2004-05-06 Pratt & Whitney Canada Corp. Plate-forme de pales refroidie passivement
EP1452693A2 (fr) * 2003-02-27 2004-09-01 General Electric Company Anneau de guidage pour une turbine
EP1512835A2 (fr) 2003-09-02 2005-03-09 General Electric Company Aube de rotor et moteur à turbine à gaz comprenant un ensemble rotor correspondant
US20050135936A1 (en) 2003-12-17 2005-06-23 Anthony Cherolis Turbine blade with trailing edge platform undercut
EP1561900A2 (fr) * 2004-02-03 2005-08-10 United Technologies Corporation Circuit de refroidissement pour plate-forme d'aube de turbine
EP1582697A1 (fr) * 2004-03-30 2005-10-05 United Technologies Corporation Injection du air de refroidissement pour turbines
CA2528049A1 (fr) * 2004-12-13 2006-06-13 Pratt & Whitney Canada Corp. Refroidissement par impact de jet de plate-forme de profil aerodynamique
DE102006004437A1 (de) * 2005-05-27 2006-11-30 Mitsubishi Heavy Industries, Ltd. Plattform einer Laufschaufel einer Gasturbine, Verfahren zur Herstellung einer Laufschaufel, Dichtungsplatte und Gasturbine
US20070009359A1 (en) 2005-02-17 2007-01-11 United Technologies Corporation Industrial gas turbine blade assembly
EP1795703A2 (fr) * 2005-12-08 2007-06-13 General Electric Company Amortisseur pour une plateforme refroidie d'aube de turbine

Patent Citations (18)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
DE1247072B (de) 1962-12-05 1967-08-10 Gen Motors Corp Hohlschaufel, insbesondere fuer Gasturbinen
DE2639511A1 (de) 1975-09-08 1977-03-17 Gen Electric Kuehlluftleckstromausnutzung
FR2417639A1 (fr) * 1976-05-14 1979-09-14 Rolls Royce Dispositif refrigerant pour aube de distributeur de moteur a turbine a gaz
DE3248161A1 (de) 1981-12-28 1983-07-07 United Technologies Corp., 06101 Hartford, Conn. Kuehlbare schaufel
GB2244520A (en) * 1990-05-31 1991-12-04 Gen Electric Nozzle assembly for a gas turbine engine
US5201847A (en) * 1991-11-21 1993-04-13 Westinghouse Electric Corp. Shroud design
US5954475A (en) * 1996-01-08 1999-09-21 Mitsubishi Jukogyo Kabushiki Kaisha Gas turbine stationary blade
EP1380726A2 (fr) * 2002-07-10 2004-01-14 Mitsubishi Heavy Industries, Ltd. Aube statorique pour turbine à gaz et turbine à gaz comprenant cet élément
WO2004038179A1 (fr) * 2002-10-24 2004-05-06 Pratt & Whitney Canada Corp. Plate-forme de pales refroidie passivement
EP1452693A2 (fr) * 2003-02-27 2004-09-01 General Electric Company Anneau de guidage pour une turbine
EP1512835A2 (fr) 2003-09-02 2005-03-09 General Electric Company Aube de rotor et moteur à turbine à gaz comprenant un ensemble rotor correspondant
US20050135936A1 (en) 2003-12-17 2005-06-23 Anthony Cherolis Turbine blade with trailing edge platform undercut
EP1561900A2 (fr) * 2004-02-03 2005-08-10 United Technologies Corporation Circuit de refroidissement pour plate-forme d'aube de turbine
EP1582697A1 (fr) * 2004-03-30 2005-10-05 United Technologies Corporation Injection du air de refroidissement pour turbines
CA2528049A1 (fr) * 2004-12-13 2006-06-13 Pratt & Whitney Canada Corp. Refroidissement par impact de jet de plate-forme de profil aerodynamique
US20070009359A1 (en) 2005-02-17 2007-01-11 United Technologies Corporation Industrial gas turbine blade assembly
DE102006004437A1 (de) * 2005-05-27 2006-11-30 Mitsubishi Heavy Industries, Ltd. Plattform einer Laufschaufel einer Gasturbine, Verfahren zur Herstellung einer Laufschaufel, Dichtungsplatte und Gasturbine
EP1795703A2 (fr) * 2005-12-08 2007-06-13 General Electric Company Amortisseur pour une plateforme refroidie d'aube de turbine

Cited By (18)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
WO2011029420A1 (fr) * 2009-09-10 2011-03-17 Mtu Aero Engines Gmbh Dispositif de déflexion pour un écoulement de fuite dans une turbine à gaz et turbine à gaz correspondante
CN102953772A (zh) * 2011-08-08 2013-03-06 通用电气公司 用于控制涡轮机中的流的***和方法
DE102013200497A1 (de) * 2013-01-15 2014-07-17 Siemens Aktiengesellschaft Drosseleinheit sowie Rotor und Gasturbine
FR3001492A1 (fr) * 2013-01-25 2014-08-01 Snecma Stator de turbomachine avec controle passif de la purge
US10287911B2 (en) 2014-11-25 2019-05-14 Safran Aircraft Engines Turbine engine rotor shaft comprising an improved heat exchange surface
FR3028883A1 (fr) * 2014-11-25 2016-05-27 Snecma Arbre de rotor de turbomachine comportant une surface d'echange thermique perfectionnee
GB2534016A (en) * 2014-11-25 2016-07-13 Snecma Turbine engine rotor shaft comprising an improved heat exchange surface
GB2534016B (en) * 2014-11-25 2020-12-02 Snecma Turbine engine rotor shaft comprising an improved heat exchange surface
US10619484B2 (en) 2015-01-22 2020-04-14 General Electric Company Turbine bucket cooling
US10196915B2 (en) 2015-06-01 2019-02-05 United Technologies Corporation Trailing edge platform seals
EP3101236A1 (fr) * 2015-06-01 2016-12-07 United Technologies Corporation Joints de plate-forme de bord de fuite
EP3273004A1 (fr) * 2016-07-22 2018-01-24 General Electric Company Refroidissement d'aube de turbine
EP3521571A1 (fr) * 2018-01-31 2019-08-07 United Technologies Corporation Refroidissement par impact de plate-forme des aubes de turbine
US10526917B2 (en) 2018-01-31 2020-01-07 United Technologies Corporation Platform lip impingement features
EP3521570A1 (fr) * 2018-02-05 2019-08-07 United Technologies Corporation Composant et moteur à turbine à gaz associé
US10822987B1 (en) 2019-04-16 2020-11-03 Pratt & Whitney Canada Corp. Turbine stator outer shroud cooling fins
FR3107919A1 (fr) * 2020-03-03 2021-09-10 Safran Aircraft Engines Aube creuse de turbomachine et plateforme inter-aubes équipées de saillies perturbatrices de flux de refroidissement
FR3107920A1 (fr) * 2020-03-03 2021-09-10 Safran Aircraft Engines Aube creuse de turbomachine et plateforme inter-aubes équipées de saillies perturbatrices de flux de refroidissement

Also Published As

Publication number Publication date
WO2009106464A1 (fr) 2009-09-03

Similar Documents

Publication Publication Date Title
EP2093381A1 (fr) Aube rotorique ou statorique de turbine à plateforme refroidie
CN104564350B (zh) 用于冷却燃气涡轮的热气体路径中的构件的布置
US7704045B1 (en) Turbine blade with blade tip cooling notches
JP5269223B2 (ja) タービン翼
US6607355B2 (en) Turbine airfoil with enhanced heat transfer
JP4063938B2 (ja) ガスタービンエンジンの動翼の冷却通路の乱流器構造
US7510367B2 (en) Turbine airfoil with endwall horseshoe cooling slot
US8192146B2 (en) Turbine blade dual channel cooling system
CA2480393C (fr) Refroidissement par impact de jet de pales ou d'ailettes de turbine a gaz
EP1327747B1 (fr) Bord de fuite d'aube de turbine refroidi par impact
US7296972B2 (en) Turbine airfoil with counter-flow serpentine channels
EP1106781A1 (fr) Aube refroidie de rotor ou de stator de turbomachine
US20080085190A1 (en) Turbine airfoil with submerged endwall cooling channel
US7985050B1 (en) Turbine blade with trailing edge cooling
US8167560B2 (en) Turbine airfoil with an internal cooling system having enhanced vortex forming turbulators
JP4929097B2 (ja) ガスタービン翼
JP2010509532A5 (fr)
US20080166240A1 (en) Advanced cooling method for combustion turbine airfoil fillets
EP1728970B1 (fr) Système de refroidissement d'une aube de turbine
EP3184743B1 (fr) Aileron de turbine avec circuit de refroidissement de bord de fuite
EP2886797B1 (fr) Une aube rotorique ou statorique d'une turbine à gaz qui est creuse et refroidi dans laquelle les canaux de refroidissement comportent des goupilles qui sont connectée par des nervures
US8708645B1 (en) Turbine rotor blade with multi-vortex tip cooling channels
US20120100008A1 (en) Annular flow channel section for a turbomachine
US8444375B2 (en) Cooled blade for a gas turbine, method for producing such a blade, and gas turbine having such a blade
US20110223036A1 (en) Blade for a gas turbine

Legal Events

Date Code Title Description
PUAI Public reference made under article 153(3) epc to a published international application that has entered the european phase

Free format text: ORIGINAL CODE: 0009012

AK Designated contracting states

Kind code of ref document: A1

Designated state(s): AT BE BG CH CY CZ DE DK EE ES FI FR GB GR HR HU IE IS IT LI LT LU LV MC MT NL NO PL PT RO SE SI SK TR

AX Request for extension of the european patent

Extension state: AL BA MK RS

AKX Designation fees paid
REG Reference to a national code

Ref country code: DE

Ref legal event code: 8566

STAA Information on the status of an ep patent application or granted ep patent

Free format text: STATUS: THE APPLICATION IS DEEMED TO BE WITHDRAWN

18D Application deemed to be withdrawn

Effective date: 20100227