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WO2009121716A1 - Aube pour turbine à gaz - Google Patents

Aube pour turbine à gaz Download PDF

Info

Publication number
WO2009121716A1
WO2009121716A1 PCT/EP2009/053116 EP2009053116W WO2009121716A1 WO 2009121716 A1 WO2009121716 A1 WO 2009121716A1 EP 2009053116 W EP2009053116 W EP 2009053116W WO 2009121716 A1 WO2009121716 A1 WO 2009121716A1
Authority
WO
WIPO (PCT)
Prior art keywords
blade
gas turbine
core
turbine
cooling
Prior art date
Application number
PCT/EP2009/053116
Other languages
German (de)
English (en)
Inventor
Beat Von Arx
Roland DÜCKERSHOFF
Brian Kenneth Wardle
Original Assignee
Alstom Technology Ltd
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 Technology Ltd filed Critical Alstom Technology Ltd
Priority to JP2011502326A priority Critical patent/JP2011516269A/ja
Priority to EP09726863A priority patent/EP2257399A1/fr
Publication of WO2009121716A1 publication Critical patent/WO2009121716A1/fr
Priority to US12/893,276 priority patent/US20110058957A1/en

Links

Classifications

    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B22—CASTING; POWDER METALLURGY
    • B22D—CASTING OF METALS; CASTING OF OTHER SUBSTANCES BY THE SAME PROCESSES OR DEVICES
    • B22D31/00—Cutting-off surplus material, e.g. gates; Cleaning and working on castings
    • B22D31/002—Cleaning, working on castings
    • 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
    • F01D5/147—Construction, i.e. structural features, e.g. of weight-saving hollow 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
    • 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
    • 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
    • F05D2230/00—Manufacture
    • F05D2230/20—Manufacture essentially without removing material
    • F05D2230/21—Manufacture essentially without removing material by casting
    • 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
    • F05D2230/00—Manufacture
    • F05D2230/60—Assembly methods

Definitions

  • the present invention relates to the field of gas turbine technology. It relates to a blade for a gas turbine according to the preamble of claim 1.
  • Such a gas turbine which has become known in the art as GT24 / 26, for example, from an article by Joos, F. et al., "Field Experience of the Sequential Combustion System for the ABB GT24 / GT26 gas turbine family, IG TI / ASME 98-GT-220, 1998 Sweden. 1 shows the basic structure of such a gas turbine, wherein the local Fig. 1 in the present application as Fig. 1 is reproduced. Furthermore, such a gas turbine is known from EP-B1 -0 620 362.
  • FIG. 1 shows a gas turbine 10 with sequential combustion, in which along an axis 19 a compressor 11, a first combustion chamber 14, a high-pressure turbine (HDT) 15, a second combustion chamber 17 and a low-pressure turbine (NDT) 18 are arranged.
  • the compressor 1 1 and the two turbines 15, 18 are part of a rotor which rotates about the axis 19.
  • the compressor 1 1 sucks in air and compresses it.
  • the compressed air flows into a plenum, and from there into premix burners, where this air is mixed with at least one fuel, fuel supplied at least via the fuel feed 12.
  • premix burners are fundamentally apparent from EP-A1-0 321 809 or EP-A2-0 704 657.
  • the compressed air flows into the premix burners, where the mixing, as stated above, takes place with at least one fuel.
  • This fuel / air mixture then flows into the first combustion chamber 14, into which this mixture passes to form a stable flame front for combustion.
  • the hot gas provided in this way is partially expanded in the subsequent high-pressure turbine 15 under working power and then flows into the second combustion chamber 17, where a further fuel supply 16 takes place. Due to the high temperatures, which still has the hot gas partially released in the high-pressure turbine 15, combustion takes place in the second combustion chamber 17, which combustion is based on autoignition.
  • the hot gas reheated in the second combustion chamber 17 is then expanded in a multistage low-pressure turbine 18.
  • the low-pressure turbine 18 comprises in the flow direction arranged one behind the other a plurality of rows of blades and vanes, which are arranged alternately.
  • the guide vanes of the third row of guide vanes in the direction of flow are designated in FIG. 1 by the reference numeral 20 '.
  • the guide vanes are provided in their interior with a mostly serpentine manner between the ends of the airfoil guided back and forth cooling channel through which a cooling medium, usually cooling air, flows. This also applies to all thermally highly loaded blades.
  • a casting method is predominantly used in which a casting core is used to form the cooling channel.
  • the casting core projects out of the blade at one or both ends and, after completion of the casting process, leaves one or more core exits correspondingly which must later be closed.
  • a method for closing such openings is described for example in the document US-B2-6, 837,417.
  • the opening in the blade is closed with a sintered cap, which connects flush neither to the inside nor on the outside of the respective wall surface. This leads to uneven, stepped surfaces, which the Prevent flow of the medium used for cooling and thus affect the effectiveness of the cooling, sometimes even questioned.
  • the invention aims to remedy this situation. It is therefore an object of the invention to provide a blade of the type mentioned, which avoids the disadvantages of known blades and is characterized by an optimized, undisturbed flow of the cooling medium in the blade.
  • the object is solved by the entirety of the features of claim 1.
  • Essential for the invention is that the closure elements are designed and inserted into the core exits, that they connect flush to the wall surface of the cooling channel. As a result, a negative influence on the flow of the cooling medium is reliably avoided by the closure elements.
  • closure elements are designed as prefabricated sealing plugs. These can be easily inserted into the core exits and fixed there quickly and securely. This is preferably done by the closure elements or sealing plugs are soldered hard into the core exits.
  • the closure element or the closure stopper can be positioned in a particularly simple manner if connecting surfaces are formed in the core exits on which the closure elements or sealing plugs rest.
  • the closure elements or sealing plugs are inserted into the core exits in such a way that they are flush with the outer surfaces of the platforms. This results in aerodynamic advantages also in the outer space of the blade.
  • the blade according to the invention is advantageously used in a gas turbine.
  • the gas turbine can be a gas turbine with sequential combustion, which has a first combustion chamber with a downstream high-pressure turbine and a second combustion chamber with a downstream low-pressure turbine, wherein the blades are arranged both in the low-pressure turbine or in the high-pressure turbine.
  • the low-pressure turbine in such a gas turbine in the flow direction behind one another several rows of guide and moving blades.
  • Fig. 1 shows the basic structure of a gas turbine with sequential
  • FIG. 2 is a perspective side view of a vane
  • Fig. 3 is a top plan view of the outer platform with a first
  • Fig. 4 shows the section through the closed core exit in the plane
  • Fig. 5 in plan view from below the inner platform with a second
  • Fig. 2 is a perspective side view of a vane, which can be used for example in the low-pressure turbine of a gas turbine with sequential combustion of Fig. 1, and is suitable for the realization of the invention.
  • the guide vane 20 used here comprises an aerofoil blade 22 which curved in space and extends in the longitudinal direction (in the radial direction of the gas turbine) between a vane head 23 and a cover plate 21 and extends in the direction of the hot gas stream 30 from a front edge 27 to a trailing edge 28. Between the two edges 27 and 28, the blade 22 is outwardly bounded by a pressure side 31 (in Fig. 2 facing the viewer) and a (opposite) suction side.
  • the vane 20 is secured by means of the formed on the top of the cover plate 21 hook-shaped fastening elements 24 and 25 on the turbine housing, while it rests sealingly with the blade head 23 on the rotor.
  • a serpentine manner between the platforms 21, 23 reciprocating cooling channel (39 in Fig. 4, 6) is provided for cooling the blade 20, as shown for example in the document WO-A1 - 2006029983.
  • a core is necessary, which in the present example in the platforms 21 and 23 leaves the core exits 40 in the cover plate (FIGS. 3, 4) or 41 in the blade head (FIGS. 5, 6) ,
  • the core exits 40, 41 are formed as shown in FIG. 4 and FIG. 6 and closed with corresponding plugs 32 and 36, that the outer surfaces of the sealing plug 32, 26 at least there flush with the wall surfaces of the environment, where the wall surfaces with the flowing cooling medium are acted upon. This is the case above all in the cooling channel 39, through which the cooling medium is conducted in the interior of the blade.
  • annular attachment surface 33 is provided in the core exit by a diameter step, on which the closure stopper 32 is seated with a corresponding shoulder (FIG. 4).
  • the sealing plug 32 is dimensioned and shaped such that after its insertion into the core outlet 40 both the outer surface of the cover plate 21 is continuous and the surface 35 of the inner wall of the cooling channel 39.
  • the sealing plug 32 is preferably by means of a brazed joint 34 in the core exit 40 fixed.
  • connecting surfaces 37 are provided on opposite sides at a predetermined depth, onto which the sealing plug 36 inserted into the core exit 41 and adapted in the edge contour is seated (FIG. 6). Again, the sealing plug 36 is fixed by means of brazed joints 38 in the core exit 41 and connects flush with the surrounding surface.
  • the invention which can be used in principle in all cooled blades of turbines, the disturbing influence of the closure elements is minimized to the flow of the cooling medium. As a result, the walls of the blade are optimally cooled, which leads to an extension of the blade life.
  • a preferred use of the inventive blade is to be found in large stationary gas turbines, for example, in gas turbines with sequential combustion, which have become known in the art under the name GT24 / 26. In the latter gas turbine, the preferred Use of such a blade in the low-pressure turbine find. For other gas turbine types, such a blade can also be used.

Landscapes

  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Architecture (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
  • Molds, Cores, And Manufacturing Methods Thereof (AREA)

Abstract

La présente invention concerne une aube (20) pour turbine à gaz, en particulier la turbine à basse pression d’une turbine à gaz à combustion séquentielle. Ladite aube est fabriquée selon un procédé de coulée et comprend un aubage s’étendant dans la direction radiale entre une plate-forme intérieure et une plate-forme extérieure (21). A l’intérieur de cet aubage s’étend un canal de refroidissement qui passe devant les plate-formes (21) et à travers lequel s’écoule un milieu réfrigérant, en particulier un air de refroidissement, destiné à refroidir les aubes (20). Des sorties de noyau (40) provenant de l’utilisation d’un noyau de coulée, qui relient le canal de refroidissement à l’extérieur et qui sont fermées par un élément de fermeture (32), sont présentes dans les plate-formes extérieure et/ou intérieure (21). Avec une telle aube (20), un refroidissement optimal est garanti par la conception des éléments de fermeture (32) et leur insertion dans les sorties de noyau (40) de telle sorte qu’ils jouxtent clairement la surface de paroi du canal de refroidissement.
PCT/EP2009/053116 2008-03-31 2009-03-17 Aube pour turbine à gaz WO2009121716A1 (fr)

Priority Applications (3)

Application Number Priority Date Filing Date Title
JP2011502326A JP2011516269A (ja) 2008-03-31 2009-03-17 ガスタービン用ブレード
EP09726863A EP2257399A1 (fr) 2008-03-31 2009-03-17 Aube pour turbine à gaz
US12/893,276 US20110058957A1 (en) 2008-03-31 2010-09-29 Blade for a gas turbine

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
CH4702008 2008-03-31
CH00470/08 2008-03-31

Related Child Applications (1)

Application Number Title Priority Date Filing Date
US12/893,276 Continuation US20110058957A1 (en) 2008-03-31 2010-09-29 Blade for a gas turbine

Publications (1)

Publication Number Publication Date
WO2009121716A1 true WO2009121716A1 (fr) 2009-10-08

Family

ID=39592112

Family Applications (1)

Application Number Title Priority Date Filing Date
PCT/EP2009/053116 WO2009121716A1 (fr) 2008-03-31 2009-03-17 Aube pour turbine à gaz

Country Status (4)

Country Link
US (1) US20110058957A1 (fr)
EP (1) EP2257399A1 (fr)
JP (1) JP2011516269A (fr)
WO (1) WO2009121716A1 (fr)

Cited By (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP2471612A1 (fr) * 2010-12-30 2012-07-04 United Technologies Corporation Procédé et noyau de moulage permettant de former un atterrissage pour le soudage d'un déflecteur inséré dans une surface portante
DE102013214932A1 (de) * 2013-07-30 2015-02-05 MTU Aero Engines AG Verfahren zum Herstellen einer Turbomaschinenschaufel
CN105705266A (zh) * 2013-11-07 2016-06-22 西门子股份公司 具有用于翼片区段铸造的一体翼片内部型芯和围带背侧壳的陶瓷铸造型芯

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US9713838B2 (en) 2013-05-14 2017-07-25 General Electric Company Static core tie rods
US9249917B2 (en) 2013-05-14 2016-02-02 General Electric Company Active sealing member
US10765728B2 (en) 2014-04-11 2020-09-08 Cellectis Method for generating immune cells resistant to arginine and/or tryptophan depleted microenvironment
US9771816B2 (en) 2014-05-07 2017-09-26 General Electric Company Blade cooling circuit feed duct, exhaust duct, and related cooling structure
US9638045B2 (en) 2014-05-28 2017-05-02 General Electric Company Cooling structure for stationary blade
US9909436B2 (en) 2015-07-16 2018-03-06 General Electric Company Cooling structure for stationary blade
US9822653B2 (en) 2015-07-16 2017-11-21 General Electric Company Cooling structure for stationary blade
GB202213805D0 (en) * 2022-09-22 2022-11-09 Rolls Royce Plc Platform for stator vane
GB202213804D0 (en) * 2022-09-22 2022-11-09 Rolls Royce Plc Platform for stator vane

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US3626568A (en) * 1969-04-23 1971-12-14 Avco Corp Method for bonding pins into holes in a hollow turbine blade
US3982854A (en) * 1971-12-20 1976-09-28 General Electric Company Friction welded metallic turbomachinery blade element
FR2511908A1 (fr) * 1981-08-26 1983-03-04 Snecma Procede de brasage-diffusion destine aux pieces en superalliages
FR2695163A1 (fr) * 1992-09-02 1994-03-04 Snecma Aube creuse pour turbomachine et son procédé de fabrication.
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EP1422402A1 (fr) * 2001-08-29 2004-05-26 Mitsubishi Heavy Industries, Ltd. Procede permettant de fermer un orifice de travail dans la partie superieure d'une ailette de turbine a gaz

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Cited By (6)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP2471612A1 (fr) * 2010-12-30 2012-07-04 United Technologies Corporation Procédé et noyau de moulage permettant de former un atterrissage pour le soudage d'un déflecteur inséré dans une surface portante
US9403208B2 (en) 2010-12-30 2016-08-02 United Technologies Corporation Method and casting core for forming a landing for welding a baffle inserted in an airfoil
US11077494B2 (en) 2010-12-30 2021-08-03 Raytheon Technologies Corporation Method and casting core for forming a landing for welding a baffle inserted in an airfoil
US11707779B2 (en) 2010-12-30 2023-07-25 Raytheon Technologies Corporation Method and casting core for forming a landing for welding a baffle inserted in an airfoil
DE102013214932A1 (de) * 2013-07-30 2015-02-05 MTU Aero Engines AG Verfahren zum Herstellen einer Turbomaschinenschaufel
CN105705266A (zh) * 2013-11-07 2016-06-22 西门子股份公司 具有用于翼片区段铸造的一体翼片内部型芯和围带背侧壳的陶瓷铸造型芯

Also Published As

Publication number Publication date
EP2257399A1 (fr) 2010-12-08
US20110058957A1 (en) 2011-03-10
JP2011516269A (ja) 2011-05-26

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