EP2184445A1 - Support d'aubes statorique axialement segmenté d'une turbine à gaz - Google Patents

Support d'aubes statorique axialement segmenté d'une turbine à gaz Download PDF

Info

Publication number
EP2184445A1
EP2184445A1 EP08019365A EP08019365A EP2184445A1 EP 2184445 A1 EP2184445 A1 EP 2184445A1 EP 08019365 A EP08019365 A EP 08019365A EP 08019365 A EP08019365 A EP 08019365A EP 2184445 A1 EP2184445 A1 EP 2184445A1
Authority
EP
European Patent Office
Prior art keywords
turbine
gas turbine
vane carrier
axial
guide
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Withdrawn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
EP08019365A
Other languages
German (de)
English (en)
Inventor
Roderick Bryk
Sascha Dr. Dungs
Martin Hartmann
Uwe Kahlstorf
Karl Dr. Klein
Oliver Dr. Lüsebrink
Mirko Milazar
Nicolas Savilius
Oliver Dr. Schneider
Shilun Dr. Sheng
Vadim Shevchenko
Gerhard Simon
Norbert Thamm
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 EP08019365A priority Critical patent/EP2184445A1/fr
Priority to CN200980144348.5A priority patent/CN102216568B/zh
Priority to EP09824439.5A priority patent/EP2342427B1/fr
Priority to JP2011533644A priority patent/JP5596042B2/ja
Priority to PCT/EP2009/061744 priority patent/WO2010052050A1/fr
Priority to US13/127,295 priority patent/US8870526B2/en
Priority to RU2011122612/06A priority patent/RU2508450C2/ru
Priority to PL09824439T priority patent/PL2342427T3/pl
Publication of EP2184445A1 publication Critical patent/EP2184445A1/fr
Withdrawn legal-status Critical Current

Links

Images

Classifications

    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D9/00—Stators
    • F01D9/02—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles
    • F01D9/04—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles forming ring or sector
    • F01D9/041—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles forming ring or sector using blades
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D25/00—Component parts, details, or accessories, not provided for in, or of interest apart from, other groups
    • F01D25/24—Casings; Casing parts, e.g. diaphragms, casing fastenings
    • F01D25/246—Fastening of diaphragms or stator-rings

Definitions

  • the invention relates to a guide vane carrier, in particular for a gas turbine, which consists of a number of axial segments.
  • Gas turbines are used in many areas to drive generators or work machines.
  • the energy content of a fuel is used to generate a rotational movement of a turbine shaft.
  • the fuel is burned in a combustion chamber, compressed air being supplied by an air compressor.
  • the working medium produced in the combustion chamber by the combustion of the fuel, under high pressure and at high temperature, is guided via a turbine unit arranged downstream of the combustion chamber, where it relaxes to perform work.
  • a number of rotor blades which are usually combined into blade groups or rows of blades, are arranged thereon and drive the turbine shaft via a momentum transfer from the working medium.
  • For guiding the flow of the working medium in the turbine unit also commonly associated between adjacent blade rows with the turbine housing and combined into rows of guide vanes are arranged.
  • the combustion chamber of the gas turbine may be embodied as a so-called annular combustion chamber, in which a plurality of circumferentially arranged around the turbine shaft burners in a common, surrounded by a high temperature resistant surrounding wall combustion chamber space.
  • the combustion chamber is designed in its entirety as an annular structure.
  • a single combustion chamber can also be provided a plurality of combustion chambers.
  • first row of guide vanes of a turbine unit which, together with the blade row immediately downstream in the flow direction of the working medium, forms a first turbine stage of the turbine unit, which is usually followed by further turbine stages.
  • the vanes are fixed in each case via a blade root, also referred to as a platform, on a guide vane carrier of the turbine unit.
  • the guide blade carrier for securing the platforms of the guide vanes comprise an insulation segment.
  • a guide ring on the guide vane support of the turbine unit is arranged in each case.
  • Such a guide ring is spaced by a radial gap of the blade tips of the fixed at the same axial position on the turbine shaft blades of the associated blade row.
  • the guide vane carrier of the gas turbine usually made of cast steel. This is suitable to withstand the high temperatures within the gas turbine and it can thus ensure safe operation of the gas turbine.
  • the guide vanes of the gas turbine can either be fastened to a common guide vane carrier or separate axial segments are provided for each turbine stage.
  • a common guide vane carrier or separate axial segments are provided for each turbine stage.
  • the turbine vane carrier is exposed to the extremely high temperatures that require high heat resistant cast steel, but there is a temperature profile that has relatively small high temperature regions and a larger, low temperature rear region.
  • the invention is therefore based on the object to provide a guide vane, which allows a technically simpler design and more flexible adaptation to the prevailing at the vane carrier temperature profile while maintaining operational safety.
  • This object is achieved according to the invention by designing at least one axial segment as a lattice tube structure.
  • the invention is based on the consideration that a more flexible adaptation to the temperature profile within the gas turbine could occur in the area of the guide blade carrier, in particular by different materials of the individual axial segments of the guide blade carrier.
  • high temperatures occur in particular in the region of the entanglement of the guide vanes and the ring segments, since these components cause a local heat input in the region of their attachment.
  • the foremost region of the guide blade carrier is exposed to a comparatively high compressor end temperature.
  • a relatively high quality material is necessary.
  • the temperature resistance of this material is not required. These areas could consist of cheaper and less expensive material.
  • the axial segments should continue to be solid in areas of low temperature. Therefore, these axial segments should be formed as a grid tube structure, ie as interconnected, arranged in the manner of a grid structure tubes or beams.
  • the respective grid pipe structure is provided on its inner and / or outer side with a sheet metal lining.
  • a sheet-metal-clad tubular construction can replace previously provided as castings sections of the vane support by a simpler structure, without jeopardizing the operational safety of the gas turbine. At the same time a smaller amount of material is needed.
  • the respective metal cladding on cooling air holes are also easier to manufacture than the cooling air holes required for castings, whereby a finer distribution to the subsequent ring segments can be provided by increasing the number of holes with the same cross-section or flow resistance.
  • the material of the respective axial segment and / or, where appropriate, of the respective sheet-metal cladding is local to that provided during operation adapted to thermal and mechanical loads.
  • a number of axial segments are welded together.
  • the individual axial segments d. H. the individual grid pipe structures and the axial segments produced as castings ensures a dimensionally stable and secure connection.
  • all axial segments are designed as a lattice tube structure.
  • the entire vane carrier can be designed as a lattice tube structure, where appropriate, different sheet metal linings are used on the inside, segment by segment. As a result, an even simpler construction of the guide vane carrier and thus the gas turbine is possible.
  • the advantages associated with the invention are, in particular, that a technically much simpler, easier and less expensive construction of a guide blade carrier and thus the entire gas turbine is possible by the design of an axial segment of a vane carrier as a lattice tube structure.
  • more favorable materials can be used in areas with lower temperature exposure and expensive high-temperature materials remain limited to the front, hot area of the gas turbine.
  • the remaining axial segments made of castings are comparatively smaller, allowing a simpler design of the vane carrier and the entire gas turbine.
  • the grid structure is less thermally conductive than a solid casting, also finds a lower heat conduction in the axial direction, in particular from the hot areas at the compressor exit in the rear cooler areas, resulting in improved cooling of the vane support and thus a lower axial and possibly also radial thermal Expansion is achieved.
  • this design shows great potential for further development of guide vane carrier, as more flexible can be addressed to thermal and mechanical requirements.
  • At the front of the turbine vane carrier there are extremely stringent requirements for maintaining the gap to the vanes and blades to ensure turbine efficiency. With the segmentation by the grid construction, the thermal expansion behavior can be set to a much better extent than before, and thus the necessary minimum gap can be reduced.
  • FIG. 1 shows in detail a half section through a guide vane carrier 1.
  • the guide vane carrier 1 is usually conical or cylindrical in shape and consists of two segments, an upper and a lower segment, the z. B. are interconnected via flanges. Only the section through the upper segment is shown.
  • the illustrated vane carrier 1 comprises a number of axial segments 24 which are welded together to form a solid structure.
  • a number of axial segments 24 of the guide blade carrier 1 is formed as a grid tube construction 26.
  • the lattice tube constructions 26 are each provided on their inner side with a sheet metal lining 28.
  • the remaining axial segments 24 are formed as castings 30.
  • the material of the cast parts 30 and the sheet metal linings 28 is in each case adapted to the thermal conditions in their respective region in the interior of the gas turbine.
  • a complete construction of the vane support 1 made of grid pipe segments would also be possible.
  • the gas turbine 101 includes a compressor 102 for combustion air, a combustion chamber 104 and a turbine unit 106 for driving the compressor 102 and a generator, not shown, or a working machine.
  • the turbine unit 106 and the compressor 102 are arranged on a common turbine shaft 108, also referred to as a turbine runner, to which the generator or the working machine is also connected and which is rotatably mounted about its central axis 109.
  • the combustor 104 which is in the form of an annular combustor, is equipped with a number of burners 110 for combustion of a liquid or gaseous fuel.
  • the turbine unit 106 has a number of rotatable blades 112 connected to the turbine shaft 108.
  • the blades 112 are annularly disposed on the turbine shaft 108 and thus form a number of blade rows.
  • the turbine unit 106 includes a number of stationary vanes 114, which are also attached in a donut-like manner to a vane support 1 of the turbine unit 106 to form rows of vanes.
  • the blades 112 serve to drive the turbine shaft 108 by momentum transfer from the turbine unit 106 flowing through the working medium M.
  • the vanes 114 serve to guide the flow of the working medium M between two seen in the flow direction of the working medium M consecutive blade rows or blade rings.
  • a successive pair of a ring of vanes 114 or a row of vanes and a ring of blades 112 or a blade row is also referred to as a turbine stage.
  • Each vane 114 has a platform 118 which is arranged to fix the respective vane 114 to a vane support 1 of the turbine unit 106 as a wall element.
  • the platform 118 is a thermally comparatively heavily loaded component, which forms the outer boundary of a hot gas channel for the turbine unit 106 flowing through the working medium M.
  • Each blade 112 is fastened to the turbine shaft 108 in an analogous manner via a platform 119, also referred to as a blade root.
  • a guide ring 121 is respectively arranged on the guide vane carrier 16 of the turbine unit 106.
  • the outer surface of each guide ring 121 is also exposed to the hot, the turbine unit 106 flowing through the working medium M and spaced in the radial direction from the outer end of the blades lying opposite him 112 through a gap.
  • the arranged between adjacent vane rows Guide rings 121 serve in particular as cover elements, which protect the inner housing in the guide blade carrier 1 or other housing-mounting components from thermal overload by the hot working medium M flowing through the turbine 106.
  • the combustion chamber 104 is configured in the exemplary embodiment as a so-called annular combustion chamber, in which a plurality of burners 110 arranged around the turbine shaft 108 in the circumferential direction open into a common combustion chamber space.
  • the combustion chamber 104 is configured in its entirety as an annular structure, which is positioned around the turbine shaft 108 around.
  • the leftmost axial segments 24 are accordingly made of a high temperature resistant material than in the gas channel downstream areas.
  • the grid tube structure further ensures a good thermal insulation of the individual cast parts 30 from each other, as a result of which thermal deformations can be minimized.

Landscapes

  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
  • Structures Of Non-Positive Displacement Pumps (AREA)
EP08019365A 2008-11-05 2008-11-05 Support d'aubes statorique axialement segmenté d'une turbine à gaz Withdrawn EP2184445A1 (fr)

Priority Applications (8)

Application Number Priority Date Filing Date Title
EP08019365A EP2184445A1 (fr) 2008-11-05 2008-11-05 Support d'aubes statorique axialement segmenté d'une turbine à gaz
CN200980144348.5A CN102216568B (zh) 2008-11-05 2009-09-10 用于燃气轮机的轴向段的导向叶片支架
EP09824439.5A EP2342427B1 (fr) 2008-11-05 2009-09-10 Support d'aubes statorique axialement segmenté d'une turbine à gaz
JP2011533644A JP5596042B2 (ja) 2008-11-05 2009-09-10 ガスタービン用の軸方向に区分化されたガイドベーンマウント
PCT/EP2009/061744 WO2010052050A1 (fr) 2008-11-05 2009-09-10 Support d'aubes fixes, segmenté axialement, pour une turbine à gaz
US13/127,295 US8870526B2 (en) 2008-11-05 2009-09-10 Axially segmented guide vane mount for a gas turbine
RU2011122612/06A RU2508450C2 (ru) 2008-11-05 2009-09-10 Сегментированная в осевом направлении обойма направляющих лопаток для газовой турбины, а также газовая турбина и газопаровая турбинная установка с сегментированной обоймой направляющих лопаток
PL09824439T PL2342427T3 (pl) 2008-11-05 2009-09-10 Dźwigar łopatek kierujących złożony z segmentów osiowych dla turbiny gazowej

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
EP08019365A EP2184445A1 (fr) 2008-11-05 2008-11-05 Support d'aubes statorique axialement segmenté d'une turbine à gaz

Publications (1)

Publication Number Publication Date
EP2184445A1 true EP2184445A1 (fr) 2010-05-12

Family

ID=40497476

Family Applications (2)

Application Number Title Priority Date Filing Date
EP08019365A Withdrawn EP2184445A1 (fr) 2008-11-05 2008-11-05 Support d'aubes statorique axialement segmenté d'une turbine à gaz
EP09824439.5A Not-in-force EP2342427B1 (fr) 2008-11-05 2009-09-10 Support d'aubes statorique axialement segmenté d'une turbine à gaz

Family Applications After (1)

Application Number Title Priority Date Filing Date
EP09824439.5A Not-in-force EP2342427B1 (fr) 2008-11-05 2009-09-10 Support d'aubes statorique axialement segmenté d'une turbine à gaz

Country Status (7)

Country Link
US (1) US8870526B2 (fr)
EP (2) EP2184445A1 (fr)
JP (1) JP5596042B2 (fr)
CN (1) CN102216568B (fr)
PL (1) PL2342427T3 (fr)
RU (1) RU2508450C2 (fr)
WO (1) WO2010052050A1 (fr)

Families Citing this family (11)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US10018052B2 (en) 2012-12-28 2018-07-10 United Technologies Corporation Gas turbine engine component having engineered vascular structure
EP2938828A4 (fr) * 2012-12-28 2016-08-17 United Technologies Corp Composant de moteur à turbine à gaz à structure maillée vasculaire artificielle
US10094287B2 (en) 2015-02-10 2018-10-09 United Technologies Corporation Gas turbine engine component with vascular cooling scheme
US10077664B2 (en) 2015-12-07 2018-09-18 United Technologies Corporation Gas turbine engine component having engineered vascular structure
US10982672B2 (en) 2015-12-23 2021-04-20 Emerson Climate Technologies, Inc. High-strength light-weight lattice-cored additive manufactured compressor components
US10557464B2 (en) 2015-12-23 2020-02-11 Emerson Climate Technologies, Inc. Lattice-cored additive manufactured compressor components with fluid delivery features
US10634143B2 (en) 2015-12-23 2020-04-28 Emerson Climate Technologies, Inc. Thermal and sound optimized lattice-cored additive manufactured compressor components
US10221694B2 (en) 2016-02-17 2019-03-05 United Technologies Corporation Gas turbine engine component having vascular engineered lattice structure
US10774653B2 (en) 2018-12-11 2020-09-15 Raytheon Technologies Corporation Composite gas turbine engine component with lattice structure
US20210332756A1 (en) * 2020-04-24 2021-10-28 General Electric Company Methods and apparatus for gas turbine frame flow path hardware cooling
US11512611B2 (en) * 2021-02-09 2022-11-29 General Electric Company Stator apparatus for a gas turbine engine

Citations (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
GB1051244A (fr) * 1962-10-09
CH417637A (de) * 1960-09-28 1966-07-31 Licentia Gmbh Mehrstufige, axial beaufschlagte Dampf- oder Gasturbine
GB2378730A (en) * 2001-08-18 2003-02-19 Rolls Royce Plc Cooling of shroud segments of turbines
WO2005008032A1 (fr) * 2003-07-11 2005-01-27 Mtu Aero Engines Gmbh Aube de construction legere pour turbine a gaz et procede de fabrication associe

Family Cites Families (40)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
GB417637A (en) 1934-02-05 1934-10-09 Otto Dietrich Rohrleitungsbau Improvements relating to high pressure pipe joints
CH421142A (de) 1965-01-12 1966-09-30 Escher Wyss Ag Gehäuse für eine Gas- oder Dampfturbine
CH425341A (de) * 1965-07-23 1966-11-30 Bbc Brown Boveri & Cie Gasturbine mit Kühlung der Schaufelträger
CS163820B1 (fr) 1966-09-23 1975-11-07
US3742705A (en) * 1970-12-28 1973-07-03 United Aircraft Corp Thermal response shroud for rotating body
JPS541703A (en) * 1977-06-06 1979-01-08 Hitachi Ltd Diaphragm of steel plate structure
GB2053367B (en) * 1979-07-12 1983-01-26 Rolls Royce Cooled shroud for a gas turbine engine
SU1263777A1 (ru) 1984-04-12 1986-10-15 Центральный Ордена Трудового Красного Знамени Научно-Исследовательский И Проектный Институт Строительных Металлоконструкций Им.Н.П.Мельникова Сварной узел трубчатых стержней
DE3509193A1 (de) * 1985-03-14 1986-09-25 MTU Motoren- und Turbinen-Union München GmbH, 8000 München Stroemungsmaschine mit innengehaeuse
JPS62182444A (ja) * 1986-02-07 1987-08-10 Hitachi Ltd ガスタ−ビン冷却空気制御方法及び装置
US4863341A (en) * 1988-05-13 1989-09-05 Westinghouse Electric Corp. Turbine having semi-isolated inlet
US4920742A (en) * 1988-05-31 1990-05-01 General Electric Company Heat shield for gas turbine engine frame
US4826397A (en) * 1988-06-29 1989-05-02 United Technologies Corporation Stator assembly for a gas turbine engine
US5104285A (en) * 1990-10-18 1992-04-14 Westinghouse Electric Corp. Low pressure inlet ring subassembly with integral staybars
FR2679296B1 (fr) * 1991-07-17 1993-10-15 Snecma Plate-forme separee inter-aube pour disque ailete de rotor de turbomachine.
FR2685936A1 (fr) * 1992-01-08 1993-07-09 Snecma Dispositif de controle des jeux d'un carter de compresseur de turbomachine.
US5391052A (en) * 1993-11-16 1995-02-21 General Electric Co. Impingement cooling and cooling medium retrieval system for turbine shrouds and methods of operation
JPH07324601A (ja) * 1994-05-31 1995-12-12 Mitsubishi Heavy Ind Ltd 蒸気タービンの仕切板構造
GB9709086D0 (en) * 1997-05-07 1997-06-25 Rolls Royce Plc Gas turbine engine cooling apparatus
JP3564266B2 (ja) 1997-07-22 2004-09-08 三菱重工業株式会社 ガスタービン静翼の支持構造
US6179560B1 (en) * 1998-12-16 2001-01-30 United Technologies Corporation Turbomachinery module with improved maintainability
GB2348466B (en) * 1999-03-27 2003-07-09 Rolls Royce Plc A gas turbine engine and a rotor for a gas turbine engine
JP2002309906A (ja) * 2001-04-11 2002-10-23 Mitsubishi Heavy Ind Ltd 蒸気冷却型ガスタービン
JP3825279B2 (ja) 2001-06-04 2006-09-27 三菱重工業株式会社 ガスタービン
FR2829176B1 (fr) * 2001-08-30 2005-06-24 Snecma Moteurs Carter de stator de turbomachine
US6514041B1 (en) * 2001-09-12 2003-02-04 Alstom (Switzerland) Ltd Carrier for guide vane and heat shield segment
EP1306521A1 (fr) * 2001-10-24 2003-05-02 Siemens Aktiengesellschaft Ailette de rotor pour une turbine à gaz et turbine à gaz avec des ailettes de rotor
US6886343B2 (en) * 2003-01-15 2005-05-03 General Electric Company Methods and apparatus for controlling engine clearance closures
US7370467B2 (en) * 2003-07-29 2008-05-13 Pratt & Whitney Canada Corp. Turbofan case and method of making
DE102004016222A1 (de) * 2004-03-26 2005-10-06 Rolls-Royce Deutschland Ltd & Co Kg Anordnung zur selbsttätigen Laufspalteinstellung bei einer zwei- oder mehrstufigen Turbine
US7007488B2 (en) * 2004-07-06 2006-03-07 General Electric Company Modulated flow turbine nozzle
SE527732C2 (sv) 2004-10-07 2006-05-23 Volvo Aero Corp Ett hölje för omslutande av en gasturbinkomponent
US7217089B2 (en) * 2005-01-14 2007-05-15 Pratt & Whitney Canada Corp. Gas turbine engine shroud sealing arrangement
FR2891300A1 (fr) * 2005-09-23 2007-03-30 Snecma Sa Dispositif de controle de jeu dans une turbine a gaz
EP1990507B1 (fr) * 2006-03-02 2015-04-15 IHI Corporation Structure de refroidissement par contact
US7610763B2 (en) * 2006-05-09 2009-11-03 United Technologies Corporation Tailorable design configuration topologies for aircraft engine mid-turbine frames
US7798775B2 (en) * 2006-12-21 2010-09-21 General Electric Company Cantilevered nozzle with crowned flange to improve outer band low cycle fatigue
DE102008000284A1 (de) * 2007-03-02 2008-09-04 Alstom Technology Ltd. Dampfturbine
FR2923525B1 (fr) * 2007-11-13 2009-12-18 Snecma Etancheite d'un anneau de rotor dans un etage de turbine
GB2462581B (en) * 2008-06-25 2010-11-24 Rolls Royce Plc Rotor path arrangements

Patent Citations (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CH417637A (de) * 1960-09-28 1966-07-31 Licentia Gmbh Mehrstufige, axial beaufschlagte Dampf- oder Gasturbine
GB1051244A (fr) * 1962-10-09
GB2378730A (en) * 2001-08-18 2003-02-19 Rolls Royce Plc Cooling of shroud segments of turbines
WO2005008032A1 (fr) * 2003-07-11 2005-01-27 Mtu Aero Engines Gmbh Aube de construction legere pour turbine a gaz et procede de fabrication associe

Also Published As

Publication number Publication date
PL2342427T3 (pl) 2013-11-29
CN102216568A (zh) 2011-10-12
EP2342427B1 (fr) 2013-06-19
RU2508450C2 (ru) 2014-02-27
US20110268580A1 (en) 2011-11-03
JP5596042B2 (ja) 2014-09-24
JP2012507652A (ja) 2012-03-29
EP2342427A1 (fr) 2011-07-13
CN102216568B (zh) 2015-11-25
US8870526B2 (en) 2014-10-28
WO2010052050A1 (fr) 2010-05-14
RU2011122612A (ru) 2012-12-20

Similar Documents

Publication Publication Date Title
EP2342427B1 (fr) Support d'aubes statorique axialement segmenté d'une turbine à gaz
EP1443275B1 (fr) Chambre de combustion
EP1636526B1 (fr) Chambre a combustion
DE3941174C2 (de) Spitzenspalt-Einstellvorrichtung für die Turbinenrotorschaufeln eines Gasturbinentriebwerks
DE2943464A1 (de) Dichtungsvorrichtung fuer ein gasturbinentriebwerk
EP2342425B1 (fr) Turbine à gaz avec plaque de fixation entre la base d'aube et le disque
EP1947293A1 (fr) Aube directrice pour turbine à gaz
WO2001009553A1 (fr) Dispositif de refroidissement par choc
EP2411631B1 (fr) Plaque d'étanchéité et système d'aube tournante
EP2347101B1 (fr) Turbine à gaz et moteur à turbine à gaz associé
EP2206885A1 (fr) Turbine à gaz
EP1731715A1 (fr) Transition d'une chambre de combustion à une turbine
EP1744014A1 (fr) Agencement de montage des aubes d'entrée d'une turbine à gaz
EP2196628A1 (fr) Support d'aube directrice
EP2347100B1 (fr) Turbine à gaz avec insert de refroidissement
EP1429077B1 (fr) Turbine à gaz
EP1398569A1 (fr) Turbine à gaz
EP2218882A1 (fr) Système de support d'aube directrice
EP1588102B1 (fr) Element de bouclier thermique, chambre de combustion et turbine a gaz
WO2004109187A1 (fr) Element de protection thermique
EP2184449A1 (fr) Support d'aube directrice, turbine à gaz et moteur à turbine à gaz ou à vapeur avec un tel support d'aube directrice
EP2194236A1 (fr) Carter de turbine
EP2352909B1 (fr) Support d'aubes directrices
EP1329594A1 (fr) Réglage du jeu d'aubes pour une turbine à gas
EP2218880A1 (fr) Système de contrôle actif de jeu pour turbine à gaz

Legal Events

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

Free format text: ORIGINAL CODE: 0009012

AK Designated contracting states

Kind code of ref document: A1

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

AX Request for extension of the european patent

Extension state: AL BA MK RS

AKY No designation fees paid
STAA Information on the status of an ep patent application or granted ep patent

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

18D Application deemed to be withdrawn

Effective date: 20101113

REG Reference to a national code

Ref country code: DE

Ref legal event code: R108

Effective date: 20110412