EP2060743B1 - Etanchéité d'un anneau de rotor dans un étage de turbine - Google Patents

Etanchéité d'un anneau de rotor dans un étage de turbine Download PDF

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Publication number
EP2060743B1
EP2060743B1 EP08163591A EP08163591A EP2060743B1 EP 2060743 B1 EP2060743 B1 EP 2060743B1 EP 08163591 A EP08163591 A EP 08163591A EP 08163591 A EP08163591 A EP 08163591A EP 2060743 B1 EP2060743 B1 EP 2060743B1
Authority
EP
European Patent Office
Prior art keywords
sheet
ring
nozzle
annular
outer rim
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.)
Active
Application number
EP08163591A
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German (de)
English (en)
French (fr)
Other versions
EP2060743A1 (fr
Inventor
Philippe Gérard Marie Hazevis
Xavier Firmin Camille Jean Lescure
Aurélien René-Pierre Massot
Jean-Luc Soupizon
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.)
Safran Aircraft Engines SAS
Original Assignee
SNECMA SAS
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 SNECMA SAS filed Critical SNECMA SAS
Publication of EP2060743A1 publication Critical patent/EP2060743A1/fr
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Publication of EP2060743B1 publication Critical patent/EP2060743B1/fr
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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
    • F01D11/00—Preventing or minimising internal leakage of working-fluid, e.g. between stages
    • F01D11/005—Sealing means between non relatively rotating elements
    • 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
    • F05D2240/11—Shroud seal segments

Definitions

  • the present invention relates to the sealing of a rotor ring in a turbine stage of a turbomachine such as an airplane turbojet or turboprop engine as described in US Pat. EP 1538306A .
  • a turbomachine turbine comprises at least one stage comprising a distributor formed of an annular row of fixed vanes and followed by a rotor wheel mounted inside a sectorized ring.
  • the distributor comprises at its downstream end an annular flange which extends radially outwards and which carries attachment means on a casing of the turbine.
  • the downstream sectorized ring comprises an upstream cylindrical rim which is held radially on a rail of the turbine casing by means of an annular locking member with a C-shaped or U-shaped section axially engaged on the casing rail and on the cylindrical rim of the ring.
  • the cylindrical rim of the ring and the casing rail are generally thermally protected by an annular plate which is mounted between the outer rim of the distributor and the upstream end of the ring to limit the passage of gas from the vein of the ring.
  • turbine radially outwardly in the annular housing space of the rim of the ring and crankcase rail.
  • the seal is not perfect and hot gas leaks from the vein of the turbine can increase the temperature of the crankcase hooks and cause the formation of cracks or cracks may destroy the hooks.
  • vanes of the distributor generally comprise cooling air flow channels taken upstream on the compressor of the turbomachine.
  • the object of the invention is in particular to provide a simple, effective and economical answer to these problems of the prior art.
  • It relates to a turbine stage comprising sealing means between the distributor and the ring sectorized which are simple and effective to prevent the passage of gas in a radial direction between the outer rim of the distributor and the upstream end of the ring.
  • a turbine stage of a turbomachine comprising a rotor wheel mounted inside a sectorized ring carried by a turbine casing, a distributor located upstream of the wheel and formed of a row.
  • annular fixed vanes this distributor comprising at its downstream end an outer annular flange having fastening means on the turbine housing, sealing means being provided between the outer rim of the distributor and the upstream end of the ring for limiting the passage of gas radially between the outer rim of the distributor and the ring, characterized in that the sealing means comprise an annular plate which extends substantially radially between the outer rim of the distributor and the end upstream of the ring and which comprises, at its inner periphery and at its outer periphery, axial support means against a downstream face of the external rim of the distributor, the annular part e median of this sheet being axially spaced from the outer rim of the distributor and being in axial bearing against the upstream end of the ring, the sheet being resiliently biased in the
  • the sealing plate according to the invention bears axially in the direction of the upstream, at its inner periphery and at its outer periphery, on the rim of the distributor, and its median annular portion bears elastically on the upstream end. of the ring.
  • the three annular support areas of the sheet on the rim of the distributor and on the ring ensure a good seal between these elements, and thus prevent the passage of gas from the turbine duct outwards, in the annular space housing the rim of the ring and the crankcase rail, and the air leaks from this space inwards into the turbine duct.
  • the ring is in abutment at its upstream end on the middle part of the sheet, itself bearing on the rim of the distributor, which can result in a slight elastic deformation in bending of the sheet. This deformation is allowed by the axial space provided between the outer rim of the distributor and the sealing plate at the middle annular portion of this sheet.
  • This axial prestress is determined to make up the manufacturing tolerances of the various parts and, in operation, to maintain the three aforementioned supports, despite differential thermal expansions of the different parts.
  • the bending deformation of the sheet may therefore be more or less important during the different operating speeds of the turbine.
  • the sheet is fixed, by means of rivets for example, on the outer rim of the distributor.
  • the sheet is for example fixed by its inner periphery on a radially inner end portion of the outer rim of the distributor.
  • the sheet is substantially flat and is, in the mounting position, plated on the downstream face of the outer rim of the distributor, covering an annular groove of the downstream face.
  • the annular groove of the downstream face of the outer rim of the distributor makes it possible to define an axial annular space between the rim of the distributor and the median annular portion of the sheet, thus allowing the bending elastic deformation of the sheet.
  • This groove is sectorized as the distributor and can be substantially continuous 360 ° around the axis of the turbine and is closed by the sheet bearing on the outer rim of the distributor, radially inside and outside of this throat.
  • the annular plate is not flat, but curved with a concavity facing upstream. It comprises, for example, an annular portion having a U-shaped or V-shaped section whose opening is oriented upstream, this annular portion being in axial abutment against the upstream end of the ring and defining an annular space with the downstream face of the outer rim of the distributor.
  • the sealing plate is sectorized as the distributor and can extend over 360 ° around the axis of the turbine. It is preferably metallic.
  • the invention also relates to a turbomachine turbine, characterized in that it comprises at least one stage as described above.
  • the invention also relates to a turbomachine, such as an airplane turbojet or turboprop, characterized in that it comprises at least one turbine stage of the aforementioned type.
  • FIG. 1 represents a turbomachine low-pressure turbine 10 comprising four stages each comprising a distributor 12 formed of an annular row of fixed vanes 14 carried by an outer casing 16 of the turbine, and a wheel 18 located downstream of the distributor 12.
  • the wheels 18 comprise disks 20 assembled axially to each other by annular flanges 22 and carrying radial vanes 24. These wheels 18 are connected to a turbine shaft (not shown) via a drive cone 26 fixed on annular flanges 22 discs.
  • Each wheel 18 is surrounded externally with a small clearance by a ring 28 formed by sectors fixed circumferentially on the casing 16 of the turbine through locking members, as will be described in more detail in the following.
  • the distributors 12 comprise walls 30 of internal and external revolution 32, respectively, which delimit between them the annular flow vein of the gases in the turbine and between which radially extend the vanes 14.
  • the outer wall 32 of the distributor 16 of the upstream stage comprises radially outer annular rims upstream 34 and downstream 36 having axial annular tabs 38 hooking, oriented upstream and intended to be engaged in corresponding axial annular grooves 40 of the casing 16 of the turbine.
  • the vanes 14 of this distributor 12 comprise cooling air circulation channels 42 coming from a supply enclosure 44 (arrows 46) located radially outside the wall 32 of the distributor. This air is partly discharged in the vein of flow of the turbine gases through orifices (not shown) formed near the trailing edge of the vanes 14 and opening into their channels 42, and partly discharged into a chamber 52 located radially inside the wall 30 of the dispenser (arrows 54). The cooling air is taken upstream on a compressor of the turbomachine and brought into the supply chamber 44 by appropriate means.
  • the ring 28 located downstream of the distributor 12 of the upstream stage comprises at its upstream end an annular hook 56 which is applied to a corresponding cylindrical rail 58 of the casing 16, and which is held radially on this rail by an annular member 60 with a C-shaped or U-shaped section which is axially engaged from upstream on the hook 56 and the rail 5 ( figure 3 ).
  • the member 60, the hook 56 and the rail 58 are housed in an annular space 62 which extends around the ring 28 between the housing 16 and the distributor 12.
  • the member 60 bears at its upstream end on a downstream face 64 of the downstream annular rim 36 of the distributor.
  • the member 60, the housing rail 58 and the hook 56 of the ring are thermally protected by an annular plate 66 which is mounted between the upstream end of the ring 28 and the downstream face 64 of the annular flange 36 of the distributor. , to limit the passage of gas from the vein of the turbine radially outwards in the annular space 62.
  • the housing rail 58 and the hook 56 of the ring are in operation subjected to high temperatures which can cause the formation of cracks or fissures likely to destroy them.
  • ducts 68 and 70 are respectively formed in the outer wall 32 and in the outer rim 36 of the distributor, to connect the Channels 42 of the vanes to the annular space 62.
  • the ducts 68 formed in the outer wall 32 of the distributor communicate at one of their ends with a channel 42 of a blade and at the other of their ends with an annular passage 72 located radially outside the wall 32 of the distributor and defined axially by the outer annular flanges 34, 36 of the distributor.
  • the ducts 70 formed in the outer rim 36 of the distributor 12 are oblique with respect to the axis of the turbine and oriented downstream towards the outside. They open at their upstream ends in the annular passage 72, and at their downstream ends on the downstream face 64 of the outer rim 36 of the distributor.
  • annular plate 66 alone does not provide a sufficient seal between the distributor 12 and the ring 28, which results in leaks of the air injected into the annular space 62 radially inwards in the vein of the turbine.
  • the invention makes it possible to provide a simple solution to this problem by means of new sealing means.
  • the sealing means according to the invention comprise an annular plate 80 which extends radially between the outer rim 36 of the distributor and the upstream end of the ring 28 and which is axially prestressed by bearing on the upstream side, by its periphery. internal and its outer periphery, on the downstream face 64 of the flange 36 and by pressing the downstream side, at its median annular portion on the upstream end of the ring 28.
  • the annular sealing plate 80 is substantially flat and is fixed by rivets 82 on the outer flange 36 of the distributor.
  • the rivets are substantially parallel to the axis of the turbine and pass through orifices 84 formed in a radially inner end portion of the sheet 80 and corresponding orifices 86 formed in a radially inner end portion of the rim 36 of the distributor.
  • the sheet 80 completely covers an annular groove 88 of the downstream face 64 of the flange 36.
  • This groove 88 has a small axial depth, for example substantially equal to the thickness of the sheet, and has a radial dimension which is slightly smaller than that of the sheet. of the sheet 80. It is formed in the distributor areas over the entire angular extent of these sectors and can extend over 360 ° around the axis of the turbine.
  • the sheet is formed of angular sectors each fixed on a distributor sector and can extend over 360 ° around the axis of the turbine.
  • the inner periphery of the sheet 80 extends over a circumference located radially inside the groove 88 and this inner periphery bears axially against a radially inner portion of the downstream face 64 of the flange 36.
  • the outer periphery of the sheet extends over a circumference located radially outside the groove and this periphery bears axially on a radially outer portion of the downstream face 64 of the flange 36.
  • the orifices 84 and 86 for mounting the rivets 82 open at one of their ends into the annular groove 88 in the vicinity of its inner periphery, and are located radially inside the ring 28. upstream end of the ring 28 is in axial abutment on the sheet 80 in an area between the rivets and the outer periphery of the groove 88.
  • the sheet 80 is elastically prestressed by the ring which exerts a sufficient force in the axial direction upstream on the sheet so that it undergoes a slight elastic deformation in bending.
  • the axial prestressing of the sheet 80 is determined on the one hand to make up the manufacturing tolerances of the different parts, and to preserve the three annular sealing zones on the distributor flange and on the ring, despite the differential thermal expansions. parts in operation.
  • the deformation of the sheet 80 may therefore vary during an operating cycle of the turbomachine.
  • the three bearing zones provide a good seal between the turbine duct and the annular space 62 for housing the upstream flange 56 of the ring and the casing rail 58.
  • the ducts 70 of the figure 3 which provide the fluidic communication between the annular passage 72 and the annular space 62 are here replaced by axial holes 90 formed in the flange 36 of the distributor and axial grooves 92 formed in the annular tabs 38 of this flange.
  • the holes 90 open at their downstream ends radially outside the sheet 80.
  • the flange 36 of the dispenser may comprise ducts 70 similar to those of FIG. figure 3 these ducts opening at their ends downstream radially outside the sheet.
  • the plate 80 Before mounting the ring 28 on the turbine casing 16, the plate 80 can already be plated by means of the rivets on the downstream face 64 of the distributor flange. The ring 28 is then mounted on the housing rail 58 and abuts on the sheet 80 to preload axially.
  • the sheet 80 is held by the rivets and extends from upstream to downstream to the outside so that only its inner periphery is in contact with the downstream face of the flange 36 Fixing the ring on the casing then makes it possible to press the outer periphery of the sheet onto the external rim of the distributor.
  • the annular plate 80 ' is not flat but curved with a concavity oriented axially upstream.
  • the sheet 80 ' comprises in the vicinity of its outer periphery a curved annular portion V-shaped section or U whose opening is oriented upstream.
  • This sheet 80 ' is mounted in the same manner as the sheet 80 described above, and its curved portion delimits an annular space 94 with the downstream face 64 of the outer flange 36 of the distributor. It is therefore not necessary to provide an annular groove on this face 64, as is the case in the embodiment of the figure 3 .
  • the sealing plate 80, 80 ' is made of metal alloy, and has a relatively small thickness of about one to a few millimeters.
  • the sheet 80, 80 'according to the invention is associated in the example shown with a distributor whose outer rim 36 comprises means for fluid communication between the annular passage 72 and the annular space 62, this sheet could be associated to a distributor without such means.
  • the sheet could also be fixed on the distributor by other fastening means than the rivets 82. It could possibly be fixed on the upstream end of the ring 28.

Landscapes

  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
  • Gasket Seals (AREA)
EP08163591A 2007-11-13 2008-09-03 Etanchéité d'un anneau de rotor dans un étage de turbine Active EP2060743B1 (fr)

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
FR0707942A FR2923525B1 (fr) 2007-11-13 2007-11-13 Etancheite d'un anneau de rotor dans un etage de turbine

Publications (2)

Publication Number Publication Date
EP2060743A1 EP2060743A1 (fr) 2009-05-20
EP2060743B1 true EP2060743B1 (fr) 2010-12-22

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Family Applications (1)

Application Number Title Priority Date Filing Date
EP08163591A Active EP2060743B1 (fr) 2007-11-13 2008-09-03 Etanchéité d'un anneau de rotor dans un étage de turbine

Country Status (9)

Country Link
US (1) US8100644B2 (ru)
EP (1) EP2060743B1 (ru)
JP (1) JP5210804B2 (ru)
CN (1) CN101435346B (ru)
CA (1) CA2644309C (ru)
DE (1) DE602008004061D1 (ru)
ES (1) ES2356701T3 (ru)
FR (1) FR2923525B1 (ru)
RU (1) RU2476710C2 (ru)

Cited By (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP4143422B1 (fr) * 2020-04-30 2024-03-20 Safran Aircraft Engines Montage d'un anneau d'etancheite sur une turbomachine aeronautique

Families Citing this family (38)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP1965311A1 (en) 2007-03-01 2008-09-03 Research In Motion Limited System and method for transformation of syndicated content for mobile delivery
FR2923525B1 (fr) * 2007-11-13 2009-12-18 Snecma Etancheite d'un anneau de rotor dans un etage de turbine
EP2184445A1 (de) * 2008-11-05 2010-05-12 Siemens Aktiengesellschaft Axial segmentierter Leitschaufelträger für einen Gasturbine
US8277172B2 (en) * 2009-03-23 2012-10-02 General Electric Company Apparatus for turbine engine cooling air management
FR2949810B1 (fr) * 2009-09-04 2013-06-28 Turbomeca Dispositif de support d'un anneau de turbine, turbine avec un tel dispositif et turbomoteur avec une telle turbine
FR2954401B1 (fr) * 2009-12-23 2012-03-23 Turbomeca Procede de refroidissement de stators de turbines et systeme de refroidissement pour sa mise en oeuvre
DE102010005153A1 (de) 2010-01-21 2011-07-28 MTU Aero Engines GmbH, 80995 Gehäusesystem für eine Axialströmungsmaschine
US20120128472A1 (en) * 2010-11-23 2012-05-24 General Electric Company Turbomachine nozzle segment having an integrated diaphragm
US20130004306A1 (en) * 2011-06-30 2013-01-03 General Electric Company Chordal mounting arrangement for low-ductility turbine shroud
ES2731206T3 (es) * 2012-03-12 2019-11-14 MTU Aero Engines AG Turbina de gas, álabe director para la carcasa de una turbina de gas, así como procedimiento para la fabricación de un álabe director
US8961108B2 (en) * 2012-04-04 2015-02-24 United Technologies Corporation Cooling system for a turbine vane
US9771818B2 (en) * 2012-12-29 2017-09-26 United Technologies Corporation Seals for a circumferential stop ring in a turbine exhaust case
WO2014165182A1 (en) * 2013-03-13 2014-10-09 United Technologies Corporation Assembly for sealing a gap between components of a turbine engine
EP2846001B1 (de) 2013-09-06 2023-01-11 MTU Aero Engines AG Montage- und Demontageverfahren eines Gasturbinenrotors und zugehörige Werkzeug
EP2863020A1 (de) * 2013-10-16 2015-04-22 Siemens Aktiengesellschaft Turbinenschaufel, Ringsegment, zugehörige Turbinenschaufelanordnung, Stator, Rotor, Turbine und Kraftwerksanlage
RU2534671C1 (ru) * 2013-11-25 2014-12-10 Российская Федерация, от имени которой выступает Министерство промышленности и торговли Российской Федерации (Минпромторг России) Статор турбины
US10400627B2 (en) * 2015-03-31 2019-09-03 General Electric Company System for cooling a turbine engine
FR3041993B1 (fr) * 2015-10-05 2019-06-21 Safran Aircraft Engines Ensemble d'anneau de turbine avec maintien axial
ES2861200T3 (es) * 2015-12-15 2021-10-06 MTU Aero Engines AG Conexión de componentes de turbomaquinaria
CN105386797B (zh) * 2015-12-29 2017-06-16 中国航空工业集团公司沈阳发动机设计研究所 一种涡轮静子结构
DE102016115610A1 (de) 2016-08-23 2018-03-01 Rolls-Royce Deutschland Ltd & Co Kg Gasturbine und Verfahren zum Aufhängen eines Turbinen-Leitschaufelsegments einer Gasturbine
US10648362B2 (en) 2017-02-24 2020-05-12 General Electric Company Spline for a turbine engine
US10655495B2 (en) 2017-02-24 2020-05-19 General Electric Company Spline for a turbine engine
FR3066225B1 (fr) * 2017-05-12 2019-05-10 Safran Aircraft Engines Turbine pour turbomachine
US10895167B2 (en) * 2017-05-30 2021-01-19 Raytheon Technologies Corporation Metering hole geometry for cooling holes in gas turbine engine
EP3412871B1 (en) * 2017-06-09 2021-04-28 Ge Avio S.r.l. Sealing arrangement for a turbine vane assembly
CN108487940B (zh) * 2018-04-04 2024-04-02 西安交通大学 一种盘式透平的喷嘴结构
FR3080145B1 (fr) * 2018-04-17 2020-05-01 Safran Aircraft Engines Distributeur en cmc avec reprise d'effort par une pince etanche
US10982559B2 (en) * 2018-08-24 2021-04-20 General Electric Company Spline seal with cooling features for turbine engines
FR3092869B1 (fr) * 2019-02-18 2023-01-13 Safran Aircraft Engines Distributeurs de turbomachine comportant un insert de contact
FR3095830B1 (fr) * 2019-05-10 2021-05-07 Safran Aircraft Engines Module de turbomachine equipe d’un dispositif de maintien de lamelles d’etancheite
US11840930B2 (en) * 2019-05-17 2023-12-12 Rtx Corporation Component with feather seal slots for a gas turbine engine
FR3100838B1 (fr) * 2019-09-13 2021-10-01 Safran Aircraft Engines Anneau d’etancheite de turbomachine
FR3109402B1 (fr) * 2020-04-15 2022-07-15 Safran Aircraft Engines Turbine pour une turbomachine
FR3129981A1 (fr) * 2021-12-03 2023-06-09 Safran Aircraft Engines Turbine pour turbomachine
FR3129980A1 (fr) * 2021-12-03 2023-06-09 Safran Aircraft Engines Turbine pour turbomachine
CN113931872B (zh) * 2021-12-15 2022-03-18 成都中科翼能科技有限公司 一种燃气轮机压气机的双层鼓筒加强型转子结构
CN114688100B (zh) * 2022-05-31 2022-09-02 成都中科翼能科技有限公司 一种燃气涡轮发动机压气机的装配方法

Family Cites Families (42)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US4199151A (en) * 1978-08-14 1980-04-22 General Electric Company Method and apparatus for retaining seals
US4425078A (en) * 1980-07-18 1984-01-10 United Technologies Corporation Axial flexible radially stiff retaining ring for sealing in a gas turbine engine
SU1436572A1 (ru) * 1987-01-15 1992-06-23 Казанский Авиационный Институт Им.А.Н.Туполева Регулируемый сопловой аппарат турбины
SU1663202A1 (ru) * 1988-05-12 1991-07-15 Государственный научно-исследовательский институт гражданской авиации Статор турбомашины
FR2635562B1 (fr) * 1988-08-18 1993-12-24 Snecma Anneau de stator de turbine associe a un support de liaison au carter de turbine
WO1992017686A1 (en) * 1991-04-02 1992-10-15 Rolls-Royce Plc Turbine casing
US5188507A (en) * 1991-11-27 1993-02-23 General Electric Company Low-pressure turbine shroud
US5333995A (en) * 1993-08-09 1994-08-02 General Electric Company Wear shim for a turbine engine
US5562408A (en) * 1995-06-06 1996-10-08 General Electric Company Isolated turbine shroud
US5797723A (en) * 1996-11-13 1998-08-25 General Electric Company Turbine flowpath seal
US6076835A (en) * 1997-05-21 2000-06-20 Allison Advanced Development Company Interstage van seal apparatus
US6164656A (en) * 1999-01-29 2000-12-26 General Electric Company Turbine nozzle interface seal and methods
US6402466B1 (en) * 2000-05-16 2002-06-11 General Electric Company Leaf seal for gas turbine stator shrouds and a nozzle band
US6652226B2 (en) * 2001-02-09 2003-11-25 General Electric Co. Methods and apparatus for reducing seal teeth wear
US6464457B1 (en) * 2001-06-21 2002-10-15 General Electric Company Turbine leaf seal mounting with headless pins
FR2829525B1 (fr) * 2001-09-13 2004-03-12 Snecma Moteurs Assemblage de secteurs d'un distributeur de turbine a un carter
FR2835563B1 (fr) * 2002-02-07 2004-04-02 Snecma Moteurs Agencement d'accrochage de secteurs en arc de cercle de distributeur porteur d'aubes
ITMI20021219A1 (it) * 2002-06-05 2003-12-05 Nuovo Pignone Spa Dispositivo di supporto semplificato per ugelli di uno stadio di una turbina a gas
US6722846B2 (en) * 2002-07-30 2004-04-20 General Electric Company Endface gap sealing of steam turbine bucket tip static seal segments and retrofitting thereof
US6893217B2 (en) * 2002-12-20 2005-05-17 General Electric Company Methods and apparatus for assembling gas turbine nozzles
JP4269829B2 (ja) * 2003-07-04 2009-05-27 株式会社Ｉｈｉ シュラウドセグメント
US7186078B2 (en) * 2003-07-04 2007-03-06 Ishikawajima-Harima Heavy Industries Co., Ltd. Turbine shroud segment
JP4269828B2 (ja) * 2003-07-04 2009-05-27 株式会社Ｉｈｉ シュラウドセグメント
FR2858652B1 (fr) * 2003-08-06 2006-02-10 Snecma Moteurs Dispositif de controle de jeu dans une turbine a gaz
ES2316994T3 (es) * 2003-08-11 2009-04-16 Siemens Aktiengesellschaft Turbina de gas con un elemento de obturacion en la region de la corona de alabes guia o de la corona de alabes de paleta de la parte de turbina.
JP4395716B2 (ja) * 2003-09-16 2010-01-13 株式会社Ｉｈｉ シールプレート構造
FR2862338B1 (fr) * 2003-11-17 2007-07-20 Snecma Moteurs Dispositif de liaison entre un distributeur et une enceinte d'alimentation pour injecteurs de fluide de refroidissement dans une turbomachine
US7207771B2 (en) * 2004-10-15 2007-04-24 Pratt & Whitney Canada Corp. Turbine shroud segment seal
US7217089B2 (en) * 2005-01-14 2007-05-15 Pratt & Whitney Canada Corp. Gas turbine engine shroud sealing arrangement
FR2885168A1 (fr) * 2005-04-27 2006-11-03 Snecma Moteurs Sa Dispositif d'etancheite pour une enceinte d'une turbomachine, et moteur d'aeronef equipe de celui-ci
FR2899275A1 (fr) * 2006-03-30 2007-10-05 Snecma Sa Dispositif de fixation de secteurs d'anneau sur un carter de turbine d'une turbomachine
FR2899274B1 (fr) * 2006-03-30 2012-08-17 Snecma Dispositif de fixation de secteurs d'anneau autour d'une roue de turbine d'une turbomachine
FR2899281B1 (fr) * 2006-03-30 2012-08-10 Snecma Dispositif de refroidissement d'un carter de turbine d'une turbomachine
GB0619426D0 (en) * 2006-10-03 2006-11-08 Rolls Royce Plc A vane arrangement
FR2906846B1 (fr) * 2006-10-06 2008-12-26 Snecma Sa Canal de transition entre deux etages de turbine
US7665961B2 (en) * 2006-11-28 2010-02-23 United Technologies Corporation Turbine outer air seal
FR2918104B1 (fr) * 2007-06-27 2009-10-09 Snecma Sa Dispositif de refroidissement des alveoles d'un disque de rotor de turbomachine a double alimentation en air.
FR2918103B1 (fr) * 2007-06-27 2013-09-27 Snecma Dispositif de refroidissement des alveoles d'un disque de rotor de turbomachine.
FR2923525B1 (fr) * 2007-11-13 2009-12-18 Snecma Etancheite d'un anneau de rotor dans un etage de turbine
FR2925107B1 (fr) * 2007-12-14 2010-01-22 Snecma Distributeur sectorise pour une turbomachine
FR2925119A1 (fr) * 2007-12-14 2009-06-19 Snecma Sa Etancheite d'une cavite de moyeu d'un carter d'echappement dans une turbomachine
FR2928963B1 (fr) * 2008-03-19 2017-12-08 Snecma Distributeur de turbine pour une turbomachine.

Cited By (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP4143422B1 (fr) * 2020-04-30 2024-03-20 Safran Aircraft Engines Montage d'un anneau d'etancheite sur une turbomachine aeronautique

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CN101435346A (zh) 2009-05-20
RU2008144750A (ru) 2010-05-20
CN101435346B (zh) 2013-01-23
FR2923525B1 (fr) 2009-12-18
EP2060743A1 (fr) 2009-05-20
US8100644B2 (en) 2012-01-24
JP5210804B2 (ja) 2013-06-12
CA2644309C (fr) 2015-12-29
FR2923525A1 (fr) 2009-05-15
DE602008004061D1 (de) 2011-02-03
ES2356701T3 (es) 2011-04-12
US20090129917A1 (en) 2009-05-21
JP2009121461A (ja) 2009-06-04
RU2476710C2 (ru) 2013-02-27
CA2644309A1 (fr) 2009-05-13

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