EP1113145A1 - Aube pour turbine a gaz avec section de mesure sur le bord de fuite - Google Patents

Aube pour turbine a gaz avec section de mesure sur le bord de fuite Download PDF

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Publication number
EP1113145A1
EP1113145A1 EP00811043A EP00811043A EP1113145A1 EP 1113145 A1 EP1113145 A1 EP 1113145A1 EP 00811043 A EP00811043 A EP 00811043A EP 00811043 A EP00811043 A EP 00811043A EP 1113145 A1 EP1113145 A1 EP 1113145A1
Authority
EP
European Patent Office
Prior art keywords
guide element
ribs
rear edge
walls
cooling
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.)
Granted
Application number
EP00811043A
Other languages
German (de)
English (en)
Other versions
EP1113145B1 (fr
Inventor
Alexander Dr. Beeck
Jörgen Ferber
Christoph Nagler
Lothar Schneider
Klaus Semmler
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.)
General Electric Technology GmbH
Original Assignee
Alstom Power Schweiz AG
Alstom SA
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 Alstom Power Schweiz AG, Alstom SA filed Critical Alstom Power Schweiz AG
Publication of EP1113145A1 publication Critical patent/EP1113145A1/fr
Application granted granted Critical
Publication of EP1113145B1 publication Critical patent/EP1113145B1/fr
Anticipated expiration legal-status Critical
Expired - Lifetime legal-status Critical Current

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Classifications

    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B22—CASTING; POWDER METALLURGY
    • B22D—CASTING OF METALS; CASTING OF OTHER SUBSTANCES BY THE SAME PROCESSES OR DEVICES
    • B22D11/00—Continuous casting of metals, i.e. casting in indefinite lengths
    • B22D11/04—Continuous casting of metals, i.e. casting in indefinite lengths into open-ended moulds
    • B22D11/0405—Rotating moulds
    • 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
    • F05D2240/122—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

Definitions

  • the present invention relates to the field of gas turbine engines Guide elements such as guide or turbine blades. It affects one of a hot Air flow around the guide element for a gas turbine, which at least in a rear edge area, in which the air flow breaks off from the guide element at least two substantially parallel, and with ribs together there are walls connected to form inner cooling channels, and which is cooled on the inside with cooling medium flowing through the cooling channels, the cooling medium at the rear edge substantially parallel to the walls between them emerges from the guide element.
  • a gas turbine comprises a multitude of elements, which consist of hot working air be flown to. Because the working air has a temperature that for many the materials from which such flow-around components are built, in particular leads to severe signs of wear after a long period of operation it is necessary to cool many of these components.
  • the cooling can be used as internal cooling be designed in which the elements are designed as hollow profiles or simple be provided with internal cooling channels through which a cooling air flow is passed becomes.
  • film cooling it is also possible to use what is known as film cooling to provide, in which the elements are acted upon by an outside cooling air film become.
  • Modern gas turbine blades usually use a combination of the above methods, i.e. an internal convective cooling system is used, which critical points also has openings for film blowing.
  • an internal convective cooling system is used, which critical points also has openings for film blowing.
  • the amount of cooling air used must be minimized. This means, that even for large components only a small cooling air mass flow Available. To the low cooling mass flows at the same time required to realize and control efficient internal heat transfer the flow cross-sections reduced accordingly. Throttle cross sections introduced become.
  • the throttling of the cooling mass flow takes place in the area of the cast blade trailing edge, near the Cooling air outlet takes place.
  • the end of the ribs is the pressure and suction side Connect the wall, set back in the axial direction, i.e. the ribs end already inside the shovel and do not reach the rear edge.
  • Figure 1 shows a section through a guide vane according to the prior art, such as it is often used in gas turbines. It is an axial to Main axis of the turbine and cut perpendicular to the plane of the airfoil through a guide vane, as is typically immediately after the combustion chamber and in front of the first row of the gas turbine for optimal flow against the blades be used.
  • the blade is designed as a hollow profile, which bounded on the suction side by a wall 10 and on the pressure side by a further wall 11 becomes. In the inflow area, the blade is widened, walls 10 and 11 are in a curve connected to each other, and is located between the walls 10 and 11 there is a central, radially extending insert 12 around which the cooling channel leads around.
  • the guide vane 30 is only one of the two axially Direction, broken ribs interconnected walls 10 and 11 limited, cooling channels run in between. Often the central one Insert 12 completely or partially enclosed by approximately axially extending ribs. These ribs converge at the rear end of the insert (16 in Fig. 1) and from there connect the suction and pressure side bucket walls. Between The ribs form approximately axial channels in which the cooling air is guided becomes.
  • the fin bank can be interrupted by one in the radial To produce plenum 18.
  • the following ribbed bench 17 can be arranged "in line" or offset to the previous ribbed.
  • the pressure and suction walls are very short ribs or so-called pin rows connected together.
  • State of the art is well, leave these internals (ribs, pins, etc.) inside the blade ends.
  • This avoids the need for the core required for casting production has a large jump in the cross-sectional area exactly at the rear edge.
  • This strong discontinuity in the cross-sectional shape of the core leads to the manufacture to a high number of core breaks.
  • the above procedure has the considerable Disadvantage that the outlet cross section of the cooling air and thus the cooling air mass flow can only be controlled insufficiently.
  • the walls also mostly have film cooling holes 13-15, through which Cooling air can flow to the outside.
  • the invention is therefore based on the object of a hot air flow flow around the guide element of a gas turbine, which at least in a rear Edge area, at which the air flow breaks off from the guide element, at least two arranged substantially parallel, and with ribs inside each other Cooling channels forming connecting walls, and which with cooling medium flowing through the cooling channels is cooled on the inside, the Coolant at the trailing edge substantially parallel to the walls between this emerges from the guide element.
  • a first preferred embodiment of the invention is characterized in that that the throughput of cooling medium through the guide element essentially through the dimensioning of the between the ribs, here so-called throttle ribs Outlet openings is determined.
  • the better one due to the arrangement Accessibility and reworkability is particularly advantageous if the Throttling of the cooling air through the throttling ribs on the rear edge is effected, and the throttling from the outside easily by drilling or the like. can be set or measured.
  • Another embodiment of the invention is characterized in that the Thickness of the guide element at the rear edge in the range from 0.5 to 5 mm, in particular is preferably in the range of 1.0 to 2.5 mm, and that the slot thickness of the Cooling air ducts between the walls at the outlet in the range of 0.3 to 2 mm, is in particular in the range from 0.8 to 1.5 mm.
  • the guiding element is designed as a guide blade arranged in front of a turbine rotor and if air is used as the cooling medium, the ones according to the invention prove Arrangement and these dimensions as particularly advantageous.
  • the invention further comprises a method for producing one of one hot air flow around the guide element of a gas turbine, which at least in a rear edge area, in which the air flow breaks off from the guide element at least two substantially parallel, and with ribs together there are walls connected to form inner cooling channels, and which is cooled on the inside with cooling medium flowing through the cooling channels, the cooling medium at the rear edge substantially parallel to the Walls between them emerges from the guide element, which is characterized by that the guiding element is manufactured in a casting process that the rear edge area with the guide element or its walls in Flow direction extending supernatant is poured, and that after the Pour the supernatant so that at least part of the ribs arranged as a throttling ribs with the rear edge essentially flush are.
  • the casting core is shaped so that the rib geometry over the rear edge of the blade is modeled in the cast core. Only after one The rib geometry is about 0.5 to 5, preferably 1 to 3 core thicknesses hidden.
  • This method makes it easy to manufacture one according to the invention Guiding element only possible. With a normal casting process, namely the effective throttle cross-section is not simply placed directly on the trailing edge become. The sudden increase in cross-section at the outlet in the casting core leads to manufacturing to a sharp increase in core breaks. This can be done while leaving a protrusion during the casting process can be avoided.
  • a preferred embodiment of the method is characterized in that no ribs are arranged between the walls in the area of the overhang, and that the throughput of cooling medium through the finished guide element essentially by the dimensioning of the arranged between the throttle ribs Outlet openings is determined. If in the area of the protrusion on any ribs can be dispensed with in the casting process, in particular in the preferred Press casting processes ("investment casting") are largely avoided. It shows furthermore that especially if the length of the supernatant is in the range from 0.5 to 3 times as large, particularly preferably of the same size, as slot thickness of the cooling air duct between the walls, such core breaks can be avoided can do without excessive post-processing after manufacture would.
  • Figure 2 a shows a section through a guide vane directly to the rear edge bordering ribs 24 between the walls 10 and 11. It is about a figure 2 corresponding axially to the main axis of the turbine and perpendicular to Blade plane running section through a guide blade.
  • the shovel is again formed as a hollow profile, which is on the suction side of a wall 10, and is delimited on the pressure side by a further wall 11.
  • Figure 2c) shows a section along the line X-X in Figure 2a), i.e. essentially parallel to the leaf plane. Immediately adjacent to the insert 12 first ribs 16.
  • the single ones Ribs of the rows 16 and 17 advantageously have a so-called division ratio, the ratio of the radial width e normal to the plane of the sheet radial spacing f, in the range of 0.25 to 0.75.
  • Another radial plenum 19 follows, followed by so-called pins 20, i.e. as simple webs formed rows of ribs which are as uniform as possible Allow distribution of the cooling air flow at the rear edge 21.
  • the division ratio (Diameter g to radial spacing h) of the pins 20 lies in the Range from 0.25 to 0.7.
  • Such a blade is usually produced using the casting process, as a rule an investment casting process.
  • This casting process can but when making the effective throttle cross section not just straight to the Trailing edge.
  • the sudden cross-sectional expansion at the outlet in the cast core leads to a sharp increase in core breaks during manufacture. However, this can be avoided if a protrusion is left in the casting process become.
  • the cooling geometry shown in the core is based on the actual one Component limit extended.
  • Figure 2b) shows the edge area of an element extended in this way beyond the rear edge by the length b. in the In the region of the protrusion, there are advantageously no more ribs.
  • the transition from the throttle geometry does not then coincide with the core holder, rather, it first occurs within the extended component Transition from the throttle geometry to a continuous radial channel instead, which can then be used as a core holder without the risk of core breakage can.
  • This transition can be optimal in various ways depending on the procedure to be designed to hold the core, i.e. it is not imperative that the two Walls as shown in Figure 2b) simply extended evenly to the rear e.g. also a gradual protruding expansion or rejuvenation resp. Thickening of the walls in the area of the overhang is conceivable.
  • the protruding geometry is after the casting to the target length of the rear edge post-processed, i.e. removed so that the throttling points coincide with the rear edge. This can e.g. together with those that are usually necessary after the fact Post-processing such as erosion and laser drilling of the film cooling holes 13-15 happen.
  • the rear edge usually has a thickness d im Range from 0.5 to 5 mm, preferably in the range from 1.0 to 2.5 mm.
  • the slit thickness c of the cooling air duct is usually in the range from 0.3 to 2.0 mm, preferably in Range from 0.8 to 1.5 mm.
  • the protrusion b above the rear edge 0.5 to 5 times, preferably 1 to 3 times, the length a of Throttle ribs 24 amount, it is particularly advantageous if the projection b is the same as the length a of the throttle ribs.

Landscapes

  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
EP00811043A 1999-12-27 2000-11-07 Aube pour turbine a gaz avec section de mesure sur le bord de fuite Expired - Lifetime EP1113145B1 (fr)

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
DE19963349 1999-12-27
DE19963349A DE19963349A1 (de) 1999-12-27 1999-12-27 Schaufel für Gasturbinen mit Drosselquerschnitt an Hinterkante

Publications (2)

Publication Number Publication Date
EP1113145A1 true EP1113145A1 (fr) 2001-07-04
EP1113145B1 EP1113145B1 (fr) 2006-04-05

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EP00811043A Expired - Lifetime EP1113145B1 (fr) 1999-12-27 2000-11-07 Aube pour turbine a gaz avec section de mesure sur le bord de fuite

Country Status (3)

Country Link
US (1) US6481966B2 (fr)
EP (1) EP1113145B1 (fr)
DE (2) DE19963349A1 (fr)

Cited By (5)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
WO2009109462A1 (fr) * 2008-03-07 2009-09-11 Alstom Technology Ltd Pale pour turbine à gaz
EP1715139A3 (fr) * 2005-04-22 2010-04-07 United Technologies Corporation Refroidissement du bord de fuite d'une aube de turbine
WO2010086419A1 (fr) 2009-01-30 2010-08-05 Alstom Technology Ltd. Aube refroidie pour turbine à gaz
EP2584145A1 (fr) * 2011-10-20 2013-04-24 Siemens Aktiengesellschaft Pale ou aube de guidage de turbine refroidie pour turbomachine
EP2565382A3 (fr) * 2011-08-30 2015-04-22 General Electric Company Profil d'aube avec agencement de broches de refroidissement

Families Citing this family (37)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
DE50106385D1 (de) * 2001-03-26 2005-07-07 Siemens Ag Verfahren zur Herstellung einer Turbinenschaufel
US6974308B2 (en) * 2001-11-14 2005-12-13 Honeywell International, Inc. High effectiveness cooled turbine vane or blade
US6607356B2 (en) * 2002-01-11 2003-08-19 General Electric Company Crossover cooled airfoil trailing edge
US6602047B1 (en) * 2002-02-28 2003-08-05 General Electric Company Methods and apparatus for cooling gas turbine nozzles
US7014424B2 (en) * 2003-04-08 2006-03-21 United Technologies Corporation Turbine element
US6932573B2 (en) 2003-04-30 2005-08-23 Siemens Westinghouse Power Corporation Turbine blade having a vortex forming cooling system for a trailing edge
US6902372B2 (en) * 2003-09-04 2005-06-07 Siemens Westinghouse Power Corporation Cooling system for a turbine blade
US7175386B2 (en) * 2003-12-17 2007-02-13 United Technologies Corporation Airfoil with shaped trailing edge pedestals
US20050235492A1 (en) * 2004-04-22 2005-10-27 Arness Brian P Turbine airfoil trailing edge repair and methods therefor
US20080031739A1 (en) * 2006-08-01 2008-02-07 United Technologies Corporation Airfoil with customized convective cooling
US7722327B1 (en) 2007-04-03 2010-05-25 Florida Turbine Technologies, Inc. Multiple vortex cooling circuit for a thin airfoil
US8070441B1 (en) 2007-07-20 2011-12-06 Florida Turbine Technologies, Inc. Turbine airfoil with trailing edge cooling channels
US7934906B2 (en) * 2007-11-14 2011-05-03 Siemens Energy, Inc. Turbine blade tip cooling system
US10156143B2 (en) * 2007-12-06 2018-12-18 United Technologies Corporation Gas turbine engines and related systems involving air-cooled vanes
US20110135446A1 (en) * 2009-12-04 2011-06-09 United Technologies Corporation Castings, Casting Cores, and Methods
US9249675B2 (en) * 2011-08-30 2016-02-02 General Electric Company Pin-fin array
US9366144B2 (en) * 2012-03-20 2016-06-14 United Technologies Corporation Trailing edge cooling
US8951004B2 (en) * 2012-10-23 2015-02-10 Siemens Aktiengesellschaft Cooling arrangement for a gas turbine component
EP3039247B1 (fr) 2013-08-28 2020-09-30 United Technologies Corporation Système de refroidissement de nervure de socle et de croisement de surface portante de moteur à turbine à gaz
US20160298465A1 (en) * 2013-12-12 2016-10-13 United Technologies Corporation Gas turbine engine component cooling passage with asymmetrical pedestals
US20150184518A1 (en) * 2013-12-26 2015-07-02 Ching-Pang Lee Turbine airfoil cooling system with nonlinear trailing edge exit slots
WO2016160029A1 (fr) 2015-04-03 2016-10-06 Siemens Aktiengesellschaft Bord de fuite de pale de turbine à canal d'armature à faible écoulement
US10502066B2 (en) 2015-05-08 2019-12-10 United Technologies Corporation Turbine engine component including an axially aligned skin core passage interrupted by a pedestal
US10323524B2 (en) * 2015-05-08 2019-06-18 United Technologies Corporation Axial skin core cooling passage for a turbine engine component
US10508554B2 (en) 2015-10-27 2019-12-17 General Electric Company Turbine bucket having outlet path in shroud
US9885243B2 (en) 2015-10-27 2018-02-06 General Electric Company Turbine bucket having outlet path in shroud
US10156145B2 (en) * 2015-10-27 2018-12-18 General Electric Company Turbine bucket having cooling passageway
JP6671149B2 (ja) * 2015-11-05 2020-03-25 三菱日立パワーシステムズ株式会社 タービン翼及びガスタービン、タービン翼の中間加工品、タービン翼の製造方法
US10370979B2 (en) * 2015-11-23 2019-08-06 United Technologies Corporation Baffle for a component of a gas turbine engine
WO2017095438A1 (fr) * 2015-12-04 2017-06-08 Siemens Aktiengesellschaft Surface portante de turbine avec agencement de refroidissement d'un bord de fuite à sollicitation
US10598025B2 (en) * 2016-11-17 2020-03-24 United Technologies Corporation Airfoil with rods adjacent a core structure
US10641103B2 (en) 2017-01-19 2020-05-05 United Technologies Corporation Trailing edge configuration with cast slots and drilled filmholes
US10718217B2 (en) * 2017-06-14 2020-07-21 General Electric Company Engine component with cooling passages
JP7078650B2 (ja) * 2017-06-30 2022-05-31 シーメンス・エナジー・グローバル・ゲーエムベーハー・ウント・コ・カーゲー 後縁機構部を有するタービン翼および鋳造コア
US10753210B2 (en) * 2018-05-02 2020-08-25 Raytheon Technologies Corporation Airfoil having improved cooling scheme
JP6636668B1 (ja) * 2019-03-29 2020-01-29 三菱重工業株式会社 高温部品、高温部品の製造方法及び流量調節方法
CN115213379B (zh) * 2022-01-10 2023-06-20 西北工业大学 单晶叶片调控缘板杂晶的工艺筋设计方法

Citations (7)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US4286924A (en) * 1978-01-14 1981-09-01 Rolls-Royce Limited Rotor blade or stator vane for a gas turbine engine
US4292008A (en) * 1977-09-09 1981-09-29 International Harvester Company Gas turbine cooling systems
GB1605180A (en) * 1974-05-16 1983-01-26 Lls Royce Ltd Method for manufacturing a blade for a gas turbine engine
US4835958A (en) * 1978-10-26 1989-06-06 Rice Ivan G Process for directing a combustion gas stream onto rotatable blades of a gas turbine
US5243759A (en) * 1991-10-07 1993-09-14 United Technologies Corporation Method of casting to control the cooling air flow rate of the airfoil trailing edge
US5864949A (en) * 1992-10-27 1999-02-02 United Technologies Corporation Tip seal and anti-contamination for turbine blades
EP0924383A2 (fr) * 1997-12-17 1999-06-23 United Technologies Corporation Aube de turbine avec refrodissement de la racine de l'arête aval

Family Cites Families (7)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US4173120A (en) * 1977-09-09 1979-11-06 International Harvester Company Turbine nozzle and rotor cooling systems
US5368441A (en) * 1992-11-24 1994-11-29 United Technologies Corporation Turbine airfoil including diffusing trailing edge pedestals
US5669759A (en) * 1995-02-03 1997-09-23 United Technologies Corporation Turbine airfoil with enhanced cooling
EP0954679B1 (fr) * 1996-06-28 2003-01-22 United Technologies Corporation Aube pouvant etre refroidie pour moteur a turbine a gaz
US6139269A (en) * 1997-12-17 2000-10-31 United Technologies Corporation Turbine blade with multi-pass cooling and cooling air addition
US6179565B1 (en) * 1999-08-09 2001-01-30 United Technologies Corporation Coolable airfoil structure
US6234754B1 (en) * 1999-08-09 2001-05-22 United Technologies Corporation Coolable airfoil structure

Patent Citations (7)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
GB1605180A (en) * 1974-05-16 1983-01-26 Lls Royce Ltd Method for manufacturing a blade for a gas turbine engine
US4292008A (en) * 1977-09-09 1981-09-29 International Harvester Company Gas turbine cooling systems
US4286924A (en) * 1978-01-14 1981-09-01 Rolls-Royce Limited Rotor blade or stator vane for a gas turbine engine
US4835958A (en) * 1978-10-26 1989-06-06 Rice Ivan G Process for directing a combustion gas stream onto rotatable blades of a gas turbine
US5243759A (en) * 1991-10-07 1993-09-14 United Technologies Corporation Method of casting to control the cooling air flow rate of the airfoil trailing edge
US5864949A (en) * 1992-10-27 1999-02-02 United Technologies Corporation Tip seal and anti-contamination for turbine blades
EP0924383A2 (fr) * 1997-12-17 1999-06-23 United Technologies Corporation Aube de turbine avec refrodissement de la racine de l'arête aval

Cited By (10)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP1715139A3 (fr) * 2005-04-22 2010-04-07 United Technologies Corporation Refroidissement du bord de fuite d'une aube de turbine
EP2538029A1 (fr) * 2005-04-22 2012-12-26 United Technologies Corporation Refroidissement du bord de fuite d'une aube de turbine
WO2009109462A1 (fr) * 2008-03-07 2009-09-11 Alstom Technology Ltd Pale pour turbine à gaz
US8182225B2 (en) 2008-03-07 2012-05-22 Alstomtechnology Ltd Blade for a gas turbine
WO2010086419A1 (fr) 2009-01-30 2010-08-05 Alstom Technology Ltd. Aube refroidie pour turbine à gaz
US8721281B2 (en) 2009-01-30 2014-05-13 Alstom Technology Ltd. Cooled blade for a gas turbine
EP2565382A3 (fr) * 2011-08-30 2015-04-22 General Electric Company Profil d'aube avec agencement de broches de refroidissement
EP2584145A1 (fr) * 2011-10-20 2013-04-24 Siemens Aktiengesellschaft Pale ou aube de guidage de turbine refroidie pour turbomachine
WO2013056975A1 (fr) * 2011-10-20 2013-04-25 Siemens Aktiengesellschaft Pale ou aubage de guidage de turbine refroidie destiné à une turbomachine
US9896942B2 (en) 2011-10-20 2018-02-20 Siemens Aktiengesellschaft Cooled turbine guide vane or blade for a turbomachine

Also Published As

Publication number Publication date
DE19963349A1 (de) 2001-06-28
US6481966B2 (en) 2002-11-19
EP1113145B1 (fr) 2006-04-05
US20010012484A1 (en) 2001-08-09
DE50012523D1 (de) 2006-05-18

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