EP1921268A1 - Aube de turbine - Google Patents

Aube de turbine Download PDF

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Publication number
EP1921268A1
EP1921268A1 EP06023274A EP06023274A EP1921268A1 EP 1921268 A1 EP1921268 A1 EP 1921268A1 EP 06023274 A EP06023274 A EP 06023274A EP 06023274 A EP06023274 A EP 06023274A EP 1921268 A1 EP1921268 A1 EP 1921268A1
Authority
EP
European Patent Office
Prior art keywords
cooling
turbine blade
elements
channel
flow inlet
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Withdrawn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
EP06023274A
Other languages
German (de)
English (en)
Inventor
Heinz-Jürgen Dr. Gross
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 EP06023274A priority Critical patent/EP1921268A1/fr
Priority to JP2009535648A priority patent/JP2010509532A/ja
Priority to US12/513,742 priority patent/US8297926B2/en
Priority to CN200780041599.1A priority patent/CN101535602B/zh
Priority to PCT/EP2007/059935 priority patent/WO2008055737A1/fr
Priority to DE502007003044T priority patent/DE502007003044D1/de
Priority to AT07820379T priority patent/ATE459785T1/de
Priority to EP07820379A priority patent/EP2087206B1/fr
Publication of EP1921268A1 publication Critical patent/EP1921268A1/fr
Priority to JP2012046594A priority patent/JP5269223B2/ja
Withdrawn legal-status Critical Current

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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
    • 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
    • F05D2240/121—Fluid guiding means, e.g. vanes related to the leading 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
    • F05D2240/303—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
    • 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
    • F05D2260/2214—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 blade.
  • Turbine blades in particular turbine blades for gas turbines are exposed to high temperatures during operation, which quickly exceed the limit of the material stress. This applies in particular to the areas in the vicinity of the flow inlet edge.
  • it has long been known to cool turbine blades suitable, so that they have a higher temperature resistance. With turbine blades, which have a higher temperature resistance, higher energy efficiencies can be achieved in particular.
  • Cooling cooling is probably the most common type of blade cooling.
  • This type of cooling cooling air is passed through channels in the interior of the blade and uses the convective effect to dissipate the heat.
  • impingement cooling a cooling air flow impinges on the blade surface from the inside. In this way, a very good cooling effect is made possible at the point of impact, but this is limited only to the narrow area of the point of impact and the closer environment.
  • This type of cooling is therefore usually used for cooling the flow inlet edge of a turbine blade, which is exposed to high temperature loads.
  • film cooling cooling air is directed out through openings in the turbine blade from the interior of the turbine blade. This cooling air flows around the turbine blade and forms an insulating layer between the hot process gas and the blade surface.
  • the types of cooling described are suitably combined depending on the application in order to achieve the most effective blade cooling possible.
  • coolants such as turbulators, which are usually provided in the form of ribs
  • turbulators which are usually provided in the form of ribs
  • These are arranged within the cooling channels provided for the convection flow, which run in the interior of the turbine blade.
  • the incorporation of fins in the cooling channels causes the flow of cooling air in the boundary layers to be detached and entangled. Due to the forced disruption of the flow, the heat transfer can be increased in the presence of a temperature difference between the cooling channel wall and the cooling air.
  • the ribbing constantly causes the flow to form new "recovery areas" in which a substantial increase in the local heat transfer coefficient can be achieved.
  • cooling channels are often formed in turbine blades parallel to and close to the flow inlet edge, to which cooling air is supplied by further cooling channels formed in the blades.
  • the convective cooling of the flow inlet edge realized in this way is usually supplemented by impingement cooling of the inner wall of the cooling channel extending near the flow inlet edge in the case of film-sensed blades.
  • convective cooling is intensified by turbulators disposed on the inner wall of the cooling duct.
  • the invention has for its object to provide a turbine blade, which can be cooled more effectively compared to known solutions both existing and in the absence of film cooling and has a longer service life.
  • This object is achieved according to the invention with a turbine blade with at least one cooling element and a cooling channel for passing a cooling medium, in the flow of which is arranged at least one cooling element, in which the at least one cooling element is formed in a cone shape.
  • the cooling elements flowed by the cooling medium are designed in the form of pegs.
  • Pintle-shaped cooling elements cause a very strong turbulence of the cooling medium, for example in the form of cooling air, which are increased by the thus forced strong disturbance of the flow at a temperature difference between a wall of the cooling channel and the cooling medium of the heat transfer, along with a significant increase of the local heat transfer coefficients.
  • the invention provided peg-shaped design of the cooling elements which are formed during operation of the turbine blade in the cooling elements thermal stresses are kept to a minimum, so that there can be no internal cracks, in particular, the thermal stresses are significantly less than the thermal stresses that form in known cooling fins. According to the invention, therefore, the entire voltage situation is improved and it can be a significant increase in the life of the cooling elements over known solutions can be achieved, with the high life of the cooling elements and a long service life and service life of the turbine blade is connected.
  • the turbine blade according to the invention can be exposed to known gas higher temperatures compared to known solutions, even if no film cooling is provided. If film cooling is provided, even higher gas temperatures are possible. This in turn gives rise to the possibility of being able to form the turbine blade according to the invention with thinner outer walls.
  • the turbine blade has a flow inlet edge extending on one side of the turbine blade, wherein the cooling channel is delimited by a wall section relative to the flow inlet edge and the at least one cooling element extends from this wall section into the cooling channel.
  • the flow inlet edge which as a rule is subject to high thermal stress, can be cooled very effectively.
  • the peg-shaped cooling elements according to the invention which extend from the wall portion into the cooling channel, and in particular cause a strong turbulence of the cooling medium, in a present temperature difference between the wall portion and the cooling medium of the heat transfer can be significantly increased, along with a substantial increase of local heat transfer coefficients. Overall, in this way, the heat in the vicinity of the flow inlet edge can be dissipated very effectively, along with a very effective cooling of the flow inlet edge.
  • the cooling channel is preferably limited by a wall portion facing the cooling channel having a curved wall surface, wherein two or more cooling elements are provided, wherein the cooling elements have a longitudinal extent extending into the cooling channel, and wherein the two or more cooling elements are directed with their longitudinal extent to the center of the curvature of the wall surface.
  • cooling elements which are directed with their longitudinal extent to the center of the curvature of the wall surface, a very effective turbulence of the cooling elements flowing against the cooling medium can be achieved.
  • the convection cooling realized by means of the cooling elements can be very effectively combined with impingement cooling, such that the cooling medium flows in a manner onto the cooling elements in such a way that it impinges on the cooling elements, so that a very high cooling effect is achieved in the respective impingement point can be, which causes in conjunction with the provided convection cooling a very effective cooling of the turbine blade according to the invention.
  • the wall section has a wall surface facing the cooling channel, and the at least one cooling element or the two or more cooling elements extend orthogonally to the wall surface or orthogonally to the curved wall surface into the cooling channel.
  • the inventively provided extent in a direction orthogonal to the wall surface of the cooling channel causes a very effective turbulence of the cooling medium, which is accompanied by a very effective cooling, in particular the flow inlet edge, since according to the invention a directed substantially perpendicular to the longitudinal extent of the cooling elements directed flow of the cooling elements with the cooling medium can.
  • the at least one cooling element or the two or more cooling elements are integrally formed with the wall portion.
  • the cooling elements have different lengths, wherein the length of the individual cooling elements is preferably adapted to a predetermined local cooling requirement.
  • Turbine blades generally have a very inhomogeneous temperature distribution during operation, which is associated with large thermal loads acting on the turbine blades, which in particular have a detrimental effect on the service life of the turbine blade.
  • an inhomogeneous temperature distribution forming along the radial direction results for the flow inlet edge.
  • the temperature distribution for example at the flow inlet edge, can be "made uniform", since according to the invention in comparatively hot places by appropriately trained cooling elements, a correspondingly strong cooling and vice versa.
  • the turbine blade according to the invention can thus be cooled in a manner which counteracts an inhomogeneous temperature distribution, which is advantageous in particular with regard to effective cooling of the flow inlet edge.
  • the cooling capacity of each individual peg-shaped cooling element is adapted over a suitably designed length to the predetermined local cooling requirement in the vicinity of the cooling element.
  • Cooling elements in the environment of which a high cooling requirement exists, have according to the invention greater length than cooling elements in the vicinity of the cooling demand, for example, for the flow inlet edge, is less pronounced.
  • the "swirl area" and the surface to be cooled are increased, along with a significant increase in the local heat transfer coefficient.
  • the invention further relates to a turbine blade, with a flow inlet edge, a cooling channel formed in the turbine blade for carrying cooling air, which extends at least partially along the flow inlet edge, and a number of cooling elements, which are arranged in the longitudinal direction of the cooling channel in this successively fixed, wherein each individual cooling element has a cooling capacity which is adapted to a predetermined cooling requirement for the flow inlet edge in the vicinity of the cooling element, and wherein the cooling channel preferably extends continuously through the turbine blade parallel to the flow inlet edge.
  • FIG. 1 shows a sketch-like sectional view of a front section of a turbine blade 10 according to the invention, with a flat sectional surface at right angles to the flow inlet edge 12.
  • a flow is parallel to the flow inlet edge 12 to the flow inlet edge 12 extending cooling channel 14 is formed (ie, a radially extending channel 14 in axially flowed turbines), which is bounded by a wall portion 24 relative to the flow inlet edge 12.
  • peg-shaped cooling elements 18 extend into the cooling channel 14, wherein the cooling elements 18 are directed with their longitudinal extent to the center of the curvature of the wall surface 16.
  • openings 22 are formed to supply cooling air to the cooling passage 14 of other cooling passages (not shown) formed in the rear of the turbine bucket 10.
  • FIG. 2 shows a further sectional illustration of the front section of the turbine blade 10 according to the invention, with a flat sectional surface parallel to the flow inlet edge 12.
  • the cooling elements 18 formed on the curved wall surface 16 of the cooling channel 14 extend orthogonally from the curved wall surface 16 into the cooling channel 14
  • FIG. 2 shows that the length of the cooling elements 18 varies in the radial direction R. According to the invention, this counteracts the inhomogeneous temperature distribution which forms along the flow inlet edge 12 when the turbine blade 10 is used.
  • the cooling elements 18 have a greater length in the middle region than in the edge regions, since, as explained above, by increasing the length of the cooling elements 18, the local heat transfer coefficient and thus the cooling capacity of the cooling elements 18 can be increased.
  • the impingement cooling comprises the impact of cooling air emerging from the openings 22 on the arched wall surface 16 or the cooling elements 18 in order to locally enable a very good cooling effect there.
  • the cooling elements 18 are directed with their longitudinal extent to the center of the curvature of the wall surface 16, a very effective impingement cooling can be provided, with which a very effective cooling of the turbine blade 10 can be provided in conjunction with the corresponding convection cooling ,
  • the cooling passage 14 is opened on both sides of the turbine blade 10 to flow the cooling air in two directions out of the cooling passage 14.
  • a temperature harmonization of the turbine blade 10 is favored, since where cooling air is needed, cooling air is also provided, and the effect of the impingement cooling is not reduced by a cross-flow.
  • the cooling elements 18 may also be formed rib-shaped, which extend along the cooling channel 14, ie in the flow direction of the cooling air.
  • the surface of the wall surface 16 is significantly increased in order to improve the cooling of the then preferably convectively cooled turbine blade 10. It is conceivable that the height of the ribs due to the aforementioned locally different temperatures at the flow inlet edge 12 can be adapted to match.

Landscapes

  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
EP06023274A 2006-11-08 2006-11-08 Aube de turbine Withdrawn EP1921268A1 (fr)

Priority Applications (9)

Application Number Priority Date Filing Date Title
EP06023274A EP1921268A1 (fr) 2006-11-08 2006-11-08 Aube de turbine
JP2009535648A JP2010509532A (ja) 2006-11-08 2007-09-20 タービン翼
US12/513,742 US8297926B2 (en) 2006-11-08 2007-09-20 Turbine blade
CN200780041599.1A CN101535602B (zh) 2006-11-08 2007-09-20 涡轮叶片
PCT/EP2007/059935 WO2008055737A1 (fr) 2006-11-08 2007-09-20 Aube de turbine
DE502007003044T DE502007003044D1 (de) 2006-11-08 2007-09-20 Turbinenschaufel
AT07820379T ATE459785T1 (de) 2006-11-08 2007-09-20 Turbinenschaufel
EP07820379A EP2087206B1 (fr) 2006-11-08 2007-09-20 Aube de turbine
JP2012046594A JP5269223B2 (ja) 2006-11-08 2012-03-02 タービン翼

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
EP06023274A EP1921268A1 (fr) 2006-11-08 2006-11-08 Aube de turbine

Publications (1)

Publication Number Publication Date
EP1921268A1 true EP1921268A1 (fr) 2008-05-14

Family

ID=37951488

Family Applications (2)

Application Number Title Priority Date Filing Date
EP06023274A Withdrawn EP1921268A1 (fr) 2006-11-08 2006-11-08 Aube de turbine
EP07820379A Not-in-force EP2087206B1 (fr) 2006-11-08 2007-09-20 Aube de turbine

Family Applications After (1)

Application Number Title Priority Date Filing Date
EP07820379A Not-in-force EP2087206B1 (fr) 2006-11-08 2007-09-20 Aube de turbine

Country Status (7)

Country Link
US (1) US8297926B2 (fr)
EP (2) EP1921268A1 (fr)
JP (2) JP2010509532A (fr)
CN (1) CN101535602B (fr)
AT (1) ATE459785T1 (fr)
DE (1) DE502007003044D1 (fr)
WO (1) WO2008055737A1 (fr)

Cited By (8)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP2236751A3 (fr) * 2009-03-30 2012-09-19 United Technologies Corporation Aube de turbine avec bord d'attaque refroidi par jets d'air
WO2015112409A1 (fr) * 2014-01-23 2015-07-30 Siemens Aktiengesellschaft Refroidissement de chambre de bord d'attaque de profil aérodynamique à impact incliné
EP3015651A1 (fr) * 2014-10-31 2016-05-04 General Electric Company Aube de turbine refroidi comportant des tourbillons dans une chambre de refroidisement
EP3165715A1 (fr) * 2015-10-15 2017-05-10 General Electric Company Aube de turbine
US10233775B2 (en) 2014-10-31 2019-03-19 General Electric Company Engine component for a gas turbine engine
US10352177B2 (en) 2016-02-16 2019-07-16 General Electric Company Airfoil having impingement openings
US10563514B2 (en) 2014-05-29 2020-02-18 General Electric Company Fastback turbulator
CN111868352A (zh) * 2018-04-17 2020-10-30 三菱动力株式会社 涡轮叶片及燃气轮机

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US8523524B2 (en) * 2010-03-25 2013-09-03 General Electric Company Airfoil cooling hole flag region
EP2584145A1 (fr) * 2011-10-20 2013-04-24 Siemens Aktiengesellschaft Pale ou aube de guidage de turbine refroidie pour turbomachine
JP2013100765A (ja) * 2011-11-08 2013-05-23 Ihi Corp インピンジ冷却機構、タービン翼及び燃焼器
JP5834876B2 (ja) 2011-12-15 2015-12-24 株式会社Ihi インピンジ冷却機構、タービン翼及び燃焼器
EP2703601B8 (fr) * 2012-08-30 2016-09-14 General Electric Technology GmbH Aube ou ailette modulaire pour turbine à gaz et turbine à gaz avec une telle pale ou aube
US9995148B2 (en) 2012-10-04 2018-06-12 General Electric Company Method and apparatus for cooling gas turbine and rotor blades
KR101513474B1 (ko) * 2013-02-27 2015-04-23 두산중공업 주식회사 터빈 블레이드
US9850762B2 (en) 2013-03-13 2017-12-26 General Electric Company Dust mitigation for turbine blade tip turns
US10001013B2 (en) * 2014-03-06 2018-06-19 General Electric Company Turbine rotor blades with platform cooling arrangements
US10690055B2 (en) 2014-05-29 2020-06-23 General Electric Company Engine components with impingement cooling features
US10422235B2 (en) 2014-05-29 2019-09-24 General Electric Company Angled impingement inserts with cooling features
US10364684B2 (en) 2014-05-29 2019-07-30 General Electric Company Fastback vorticor pin
US9957816B2 (en) 2014-05-29 2018-05-01 General Electric Company Angled impingement insert
JP6250223B2 (ja) * 2014-07-09 2017-12-20 シーメンス アクチエンゲゼルシヤフトＳｉｅｍｅｎｓ Ａｋｔｉｅｎｇｅｓｅｌｌｓｃｈａｆｔ 内部冷却システム内のインピンジメントジェット衝突チャネルシステム
US10280785B2 (en) 2014-10-31 2019-05-07 General Electric Company Shroud assembly for a turbine engine
US20160333701A1 (en) * 2015-05-12 2016-11-17 United Technologies Corporation Airfoil impingement cavity
KR101906701B1 (ko) * 2017-01-03 2018-10-10 두산중공업 주식회사 가스터빈 블레이드
EP3396297A1 (fr) * 2017-04-28 2018-10-31 Siemens Aktiengesellschaft Dispositif de refroidissement
US10830049B2 (en) 2017-05-02 2020-11-10 Raytheon Technologies Corporation Leading edge hybrid cavities and cores for airfoils of gas turbine engine
US10907480B2 (en) * 2018-09-28 2021-02-02 Raytheon Technologies Corporation Ribbed pin fins
CN113374535A (zh) * 2021-06-28 2021-09-10 常州大学 一种格子阵列式双层冷却燃气涡轮叶片

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GB1350424A (en) * 1971-07-02 1974-04-18 Rolls Royce Cooled blade for a gas turbine engine
US5468125A (en) * 1994-12-20 1995-11-21 Alliedsignal Inc. Turbine blade with improved heat transfer surface
JPH08296403A (ja) * 1995-04-25 1996-11-12 Toshiba Corp ガスタービン空冷翼
EP0945595A2 (fr) * 1998-03-26 1999-09-29 Mitsubishi Heavy Industries, Ltd. Aube refroidie pour turbine à gaz
EP1077311A1 (fr) * 1999-08-17 2001-02-21 Siemens Aktiengesellschaft Aube refroidie de turbine à gaz
WO2004035992A1 (fr) * 2002-10-18 2004-04-29 Alstom Technology Ltd. Composant pouvant etre refroidi
EP1508746A1 (fr) * 2003-08-14 2005-02-23 Mitsubishi Heavy Industries, Ltd. Paroi d'échange de chaleur, turbine à gaz et aeronef avec une telle paroi

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JPS6163401U (fr) * 1984-06-20 1986-04-30
JPS61187501A (ja) * 1985-02-15 1986-08-21 Hitachi Ltd 流体冷却構造
JPH0663442B2 (ja) * 1989-09-04 1994-08-22 株式会社日立製作所 タービン翼
FR2678318B1 (fr) * 1991-06-25 1993-09-10 Snecma Aube refroidie de distributeur de turbine.
EP0954679B1 (fr) * 1996-06-28 2003-01-22 United Technologies Corporation Aube pouvant etre refroidie pour moteur a turbine a gaz
US5738493A (en) * 1997-01-03 1998-04-14 General Electric Company Turbulator configuration for cooling passages of an airfoil in a gas turbine engine
US6142734A (en) * 1999-04-06 2000-11-07 General Electric Company Internally grooved turbine wall
US6890153B2 (en) * 2003-04-29 2005-05-10 General Electric Company Castellated turbine airfoil
US7104757B2 (en) * 2003-07-29 2006-09-12 Siemens Aktiengesellschaft Cooled turbine blade
DE50309922D1 (de) * 2003-07-29 2008-07-10 Siemens Ag Gekühlte Turbinenschaufel

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Publication number Priority date Publication date Assignee Title
GB1350424A (en) * 1971-07-02 1974-04-18 Rolls Royce Cooled blade for a gas turbine engine
US5468125A (en) * 1994-12-20 1995-11-21 Alliedsignal Inc. Turbine blade with improved heat transfer surface
JPH08296403A (ja) * 1995-04-25 1996-11-12 Toshiba Corp ガスタービン空冷翼
EP0945595A2 (fr) * 1998-03-26 1999-09-29 Mitsubishi Heavy Industries, Ltd. Aube refroidie pour turbine à gaz
EP1077311A1 (fr) * 1999-08-17 2001-02-21 Siemens Aktiengesellschaft Aube refroidie de turbine à gaz
WO2004035992A1 (fr) * 2002-10-18 2004-04-29 Alstom Technology Ltd. Composant pouvant etre refroidi
EP1508746A1 (fr) * 2003-08-14 2005-02-23 Mitsubishi Heavy Industries, Ltd. Paroi d'échange de chaleur, turbine à gaz et aeronef avec une telle paroi

Cited By (9)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP2236751A3 (fr) * 2009-03-30 2012-09-19 United Technologies Corporation Aube de turbine avec bord d'attaque refroidi par jets d'air
US8348613B2 (en) 2009-03-30 2013-01-08 United Technologies Corporation Airflow influencing airfoil feature array
WO2015112409A1 (fr) * 2014-01-23 2015-07-30 Siemens Aktiengesellschaft Refroidissement de chambre de bord d'attaque de profil aérodynamique à impact incliné
US10563514B2 (en) 2014-05-29 2020-02-18 General Electric Company Fastback turbulator
EP3015651A1 (fr) * 2014-10-31 2016-05-04 General Electric Company Aube de turbine refroidi comportant des tourbillons dans une chambre de refroidisement
US10233775B2 (en) 2014-10-31 2019-03-19 General Electric Company Engine component for a gas turbine engine
EP3165715A1 (fr) * 2015-10-15 2017-05-10 General Electric Company Aube de turbine
US10352177B2 (en) 2016-02-16 2019-07-16 General Electric Company Airfoil having impingement openings
CN111868352A (zh) * 2018-04-17 2020-10-30 三菱动力株式会社 涡轮叶片及燃气轮机

Also Published As

Publication number Publication date
CN101535602B (zh) 2012-01-11
EP2087206B1 (fr) 2010-03-03
WO2008055737A1 (fr) 2008-05-15
US8297926B2 (en) 2012-10-30
ATE459785T1 (de) 2010-03-15
US20100143153A1 (en) 2010-06-10
CN101535602A (zh) 2009-09-16
DE502007003044D1 (de) 2010-04-15
JP5269223B2 (ja) 2013-08-21
JP2012137089A (ja) 2012-07-19
JP2010509532A (ja) 2010-03-25
EP2087206A1 (fr) 2009-08-12

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