US20120237342A1 - Turbine stage in a turbine engine - Google Patents

Turbine stage in a turbine engine Download PDF

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Publication number
US20120237342A1
US20120237342A1 US13/511,266 US201013511266A US2012237342A1 US 20120237342 A1 US20120237342 A1 US 20120237342A1 US 201013511266 A US201013511266 A US 201013511266A US 2012237342 A1 US2012237342 A1 US 2012237342A1
Authority
US
United States
Prior art keywords
casing
ring
annular
downstream
tab
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.)
Abandoned
Application number
US13/511,266
Other languages
English (en)
Inventor
Emmanuel Berche
Vincent Philippot
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
Assigned to SNECMA reassignment SNECMA ASSIGNMENT OF ASSIGNORS INTEREST (SEE DOCUMENT FOR DETAILS). Assignors: BERCHE, EMMANUEL, PHILIPPOT, VINCENT
Publication of US20120237342A1 publication Critical patent/US20120237342A1/en
Abandoned legal-status Critical Current

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Classifications

    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D11/00—Preventing or minimising internal leakage of working-fluid, e.g. between stages
    • F01D11/08—Preventing or minimising internal leakage of working-fluid, e.g. between stages for sealing space between rotor blade tips and stator
    • 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
    • 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
    • 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
    • F05D2230/64—Assembly methods using positioning or alignment devices for aligning or centring, e.g. pins
    • F05D2230/642—Assembly methods using positioning or alignment devices for aligning or centring, e.g. pins using maintaining alignment while permitting differential dilatation
    • 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
    • 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
    • F05D2240/55—Seals
    • 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
    • F05D2300/00—Materials; Properties thereof
    • F05D2300/10—Metals, alloys or intermetallic compounds
    • 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
    • F05D2300/00—Materials; Properties thereof
    • F05D2300/60—Properties or characteristics given to material by treatment or manufacturing
    • F05D2300/603—Composites; e.g. fibre-reinforced
    • Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
    • Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
    • Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
    • Y02T50/00—Aeronautics or air transport
    • Y02T50/60—Efficient propulsion technologies, e.g. for aircraft

Definitions

  • the present invention relates to a turbine stage in a turbine engine such as a turboprop or a turbojet.
  • a turbine engine essentially comprises, going from upstream to downstream: a compressor, a combustion chamber, and a turbine, the compressor feeds the combustion chamber with air under pressure, and the turbine receives the hot gas coming from the combustion chamber in order to extract energy therefrom.
  • a low pressure turbine stage comprises a nozzle constituted by an annular row of stationary vanes extending radially between two annular platforms, namely an inner platform and an outer platform, and a rotor wheel mounted downstream from the nozzle inside a sectorized ring carried by a casing surrounding the turbine stage.
  • Each ring sector carries a sealing lining on an inside face that co-operates with the outer peripheries of the blades of the rotor wheel, and it includes on an outside face means for fastening to the casing, which means are formed by upstream and downstream circumferential rims.
  • the upstream circumferential rim is engaged axially in an annular groove carried by an upstream annular tab of the casing, and the downstream circumferential rim is clamped radially against a downstream annular tab of the casing by a C-section fastener member engaged axially from downstream onto the downstream circumferential rim and the downstream annular tab.
  • An annular cavity is defined between the ring and the casing and it extends from upstream to downstream between annular tabs of the casing.
  • the upstream annular tab includes orifices feeding the cavity with air taken from a compression stage of the turbine engine.
  • each ring sector expands and deforms, taking up a curved shape that is concave in the circumferential direction, with its concave side facing outwards (negative camber phenomenon).
  • radial spaces are observed to form between the downstream annular tab of the casing and the downstream circumferential rims of the ring sectors.
  • the downstream fastening between the downstream circumferential rim of the ring and the downstream annular tab of the casing is sealed by axially prestressing the downstream annular tab against a radial face of the downstream circumferential rim that is opposite to the fastener member.
  • a particular object of the invention is to provide a solution to these problems that is simple, inexpensive, and effective, and that makes it possible to avoid the drawbacks of the prior art.
  • the invention provides a turbine stage for a turbine engine, the stage comprising a bladed wheel rotatable inside a sectorized ring made of composite material and carried by a casing, each ring sector having a downstream circumferential rim held to bear radially against an annular tab of the casing by a C-shaped fastener, the stage being characterized in that the annular tab of the casing is engaged radially in an annular groove in the downstream circumferential rim of the ring with axial clearance, when cold, that is designed to be reduced to zero in operation, when hot, and to enable leaktight axial clamping of the annular tab of the casing in the annular groove of the ring sector.
  • the downstream circumferential rim of the ring is sealed in operation by axial clamping of the upstream and downstream ends of the downstream annular tab of the casing in the annular groove as a result of the greater expansion of the casing compared with the composite ring.
  • the concave curvature of the ring and of its downstream circumferential rim is thus compensated by the axial clamping of the annular tab, thereby guaranteeing sealing of the downstream fastening of the ring.
  • the annular tab of the casing has upstream and downstream radial faces for bearing when hot against radial flanks of the groove.
  • the radial faces of the annular tab and the radial flanks of the groove conserve their radial shape, thereby ensuring annular contact between the radial faces of the ring and the radial flanks of the groove.
  • the above-mentioned axial clearance when cold is of the order of one-tenth of a millimeter.
  • annular sealing ring housed in an annular groove in the face of the annular tab that is pressed against the bottom wall of the annular groove of the ring sector.
  • the composite material is of the ceramic matrix type and the casing is made of a metal material.
  • the invention also provides a turbine engine such as an airplane turboprop or turbojet, the engine including a high pressure turbine stage of the above-described type.
  • FIG. 1 is a fragmentary diagrammatic view in axial section of a prior art turbine stage
  • FIG. 2 is a diagrammatic view in cross-section on section plane A-A shown in FIG. 1 ;
  • FIG. 3 is a fragmentary diagrammatic view in axial section of a turbine stage of the invention while cold, the section plane not passing through a fastener member;
  • FIG. 4 is a fragmentary diagrammatic view in axial section of a turbine stage of the invention while hot, on a section plane that contains a fastener member.
  • FIG. 1 shows a portion of a turbine stage 10 in a turbine engine that includes a nozzle stage having a plurality of stationary vanes arranged upstream from a rotary wheel carrying a plurality of blades and rotatable inside a ring 12 carried by an outer casing 14 .
  • the ring 12 is made up of a plurality of substantially cylindrical ring sectors that are circularly juxtaposed, end to end.
  • Each ring sector comprises a cylindrical portion 16 carrying on its inside face a sealing lining 18 of abradable material that cooperates with the outer peripheries of the blades of the rotor wheel.
  • Each ring sector includes two annular tabs, namely an upstream tab 18 and a downstream tab 20 for attaching to the casing 14 .
  • the outer end of the upstream annular tab 18 has a circumferential rim 22 extending upstream and engaged axially in an annular groove 24 facing downstream and formed in a radial annular tab 26 of the casing.
  • the outer end of the downstream annular tab 20 of the ring has a circumferential rim 28 facing downstream and clamped radially against a cylindrical portion 30 of an annular tab 32 of the casing 14 by means of a C-section fastener member 32 engaged axially on the downstream circumferential rim 28 and on the cylindrical portion 30 of the downstream annular tab 32 of the casing 14 .
  • Each downstream circumferential rim 20 of a ring sector includes at least one notch in radial alignment with a notch in the cylindrical portion 30 of the downstream annular tab 32 of the casing 14 and of width that is sufficient to enable the fastener member 32 to be engaged axially therein and to enable the ring 12 to be fastened on the casing 14 .
  • An annular cavity 34 is defined between the sectorized ring 12 and the casing 14 being defined upstream by the upstream annular tabs 18 , 26 of the ring 12 and of the casing 14 , respectively, and downstream by the downstream annular tabs 20 , 32 of the ring 12 and of the casing 14 , respectively.
  • the upstream annular tab 26 of the casing 14 has orifices 36 for passing cooling air coming from a space surrounding the combustion chamber, i.e. air that flows between the outer casing of the combustion chamber and the outer wall of the combustion chamber, which wall forms a body of revolution.
  • an annular sealing ring 38 is mounted in an annular groove 40 of the inside face of the cylindrical portion 30 .
  • This sealing ring 38 is compressed radially in the annular groove 40 and against the downstream circumferential rim 28 of the ring 12 .
  • the inside face of the cylindrical portion 30 includes a rib 42 engaged radially in an annular recess in the downstream circumferential rim 28 of the ring 12 in order to prevent the ring 12 from moving axially relative to the casing 14 .
  • each ring sector includes three slots 44 , 46 , 48 , each of which houses a sealing strip.
  • a first slot 44 is formed in the cylindrical portion 16 of the ring 12 and extends over substantially the entire length of the ring 12 , being parallel to the longitudinal axis of the ring 12 .
  • the other two slots 46 and 48 are oblique, each being formed in a respective one of the upstream and downstream annular tabs 18 and 20 of the ring.
  • the radially inner ends of the two oblique slots 46 and 48 open out into a middle portion of the longitudinal slot 44 , and their radial ends open out into the outside faces of the upstream and downstream circumferential rims 22 and 28 respectively.
  • Each sealing strip is inserted half in a slot 44 , 46 , 48 of one sector with the other half in a corresponding facing slot that is formed in a radial face of an adjacent ring sector.
  • each sector of the composite ring deforms under the effect of temperature and adopts a concave curved shape with its concave side facing outwards ( FIG. 2 ).
  • the casing 14 is also subjected to deformation and has circumferential undulations.
  • a radial space R forms between each circumferential rim 28 and the cylindrical portion 30 of a downstream tab 32 of the casing 14 , thereby leading to leaks of ventilation air from the annular cavity 34 towards the gas flow passage through the turbine.
  • the invention serves to remedy this problem and those mentioned above by forming an annular groove 50 in the outside cylindrical face of the downstream circumferential rim 52 of the ring 54 , which groove receives radially the downstream cylindrical portion 55 of the downstream annular tab 56 of the casing 14 with axial clearance j, when cold, that is designed to be reduced to zero in operation as a result of the greater expansion of the casing 14 and of its downstream annular tab 56 compared with the expansion of the ring 54 made of composite material ( FIG. 3 ).
  • the annular groove 50 has two radial annular flanks, an upstream flank 58 and a downstream flank 60 .
  • the downstream cylindrical portion 55 of the downstream annular tab 56 of the casing 14 has two radial faces, an upstream face 62 and a downstream face 64 .
  • the radial faces 62 and 64 of the downstream annular tab 56 of the casing 14 come to bear against the radial flanks 58 and 60 of the groove 50 as a result of the differential expansion between the composite ring 54 and the casing 14 , thereby ensuring that the annular tab 56 is clamped axially in the groove 50 and establishes sealing against the ventilation air flowing in the cavity 34 .
  • This axial clamping also serves to hold the ring 54 axially relative to the casing 14 .
  • the depth of the groove 50 is selected in such a manner as to be greater than the maximum radial difference R in operation between the inside face 66 of the downstream cylindrical portion 55 of the downstream tab 56 of the casing 14 and the bottom wall 68 of the groove 50 , so as to ensure continuous leakproof axial clamping when hot and so as to avoid any axial separation of the ring 54 relative to the casing 14 .
  • a ring sector is assembled by inserting the upstream circumferential rim 22 of the ring 54 in the annular groove of the upstream tab 18 of the casing 14 , and then tilting the downstream end of the ring outwards so that the cylindrical portion 55 bears against the bottom wall of the groove 50 .
  • the axial clearance j when cold, serves to make it easier to tilt the ring 54 outwards against the casing 14 .
  • An annular sealing ring 38 is housed in an annular groove 40 of the face 66 of the downstream annular tab 56 of the casing that presses against the bottom wall 68 of the groove 50 .
  • each downstream circumferential rim 52 of a ring sector includes a notch in radial alignment with a notch in the cylindrical portion of the downstream annular tab of the casing to enable the C-section fastener member 32 to be mounted axially.
  • inter-sector sealing means are similar to those of the prior art. Nevertheless, it should be observed that in the invention the sloping slot in the downstream annular tab 64 of the ring 54 opens out into the groove 50 in register with the sealing ring 38 .
  • the axial clearance, when cold, is of the order of 0.1 millimeters.
  • the ring 54 may be made of a ceramic matrix composite material that withstands well the high temperatures of the kind that exist in a high pressure turbine, and the casing 14 is made of a metal material such as Inco or steel.

Landscapes

  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
  • Gasket Seals (AREA)
US13/511,266 2009-12-18 2010-12-14 Turbine stage in a turbine engine Abandoned US20120237342A1 (en)

Applications Claiming Priority (3)

Application Number Priority Date Filing Date Title
FR09/06162 2009-12-18
FR0906162A FR2954400B1 (fr) 2009-12-18 2009-12-18 Etage de turbine dans une turbomachine
PCT/FR2010/052721 WO2011073570A1 (fr) 2009-12-18 2010-12-14 Etage de turbine dans une turbomachine

Publications (1)

Publication Number Publication Date
US20120237342A1 true US20120237342A1 (en) 2012-09-20

Family

ID=42331016

Family Applications (1)

Application Number Title Priority Date Filing Date
US13/511,266 Abandoned US20120237342A1 (en) 2009-12-18 2010-12-14 Turbine stage in a turbine engine

Country Status (8)

Country Link
US (1) US20120237342A1 (de)
EP (1) EP2513428A1 (de)
CN (1) CN102667066A (de)
BR (1) BR112012010257A2 (de)
CA (1) CA2777370A1 (de)
FR (1) FR2954400B1 (de)
RU (1) RU2012130351A (de)
WO (1) WO2011073570A1 (de)

Cited By (9)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US20150192026A1 (en) * 2013-09-06 2015-07-09 MTU Aero Engines AG Gas turbine
US20160333784A1 (en) * 2015-05-15 2016-11-17 Rolls-Royce Plc A wall cooling arrangement for a gas turbine engine
US20180037511A1 (en) * 2015-06-10 2018-02-08 Ihi Corporation Turbine
US10100659B2 (en) 2014-12-16 2018-10-16 Rolls-Royce North American Technologies Inc. Hanger system for a turbine engine component
US20190040758A1 (en) * 2015-10-05 2019-02-07 Safran Aircraft Engines Turbine ring assembly with axial retention
US10378386B2 (en) * 2015-12-18 2019-08-13 Safran Aircraft Engines Turbine ring assembly with support when cold and when hot
US10605120B2 (en) * 2016-09-27 2020-03-31 Safran Aircraft Engines Turbine ring assembly that can be set while cold
CN113898414A (zh) * 2021-12-09 2022-01-07 成都中科翼能科技有限公司 一种燃气轮机高压转子防热振动变形的补强结构
US12037926B2 (en) 2020-06-18 2024-07-16 Siemens Energy Global GmbH & Co. KG Rotor comprising a rotor component arranged between two rotor discs

Families Citing this family (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
FR3061738B1 (fr) * 2017-01-12 2019-05-31 Safran Aircraft Engines Ensemble d'anneau de turbine
US10815812B2 (en) 2017-05-12 2020-10-27 Raytheon Technologies Corporation Geometry optimized blade outer air seal for thermal loads
KR20220078706A (ko) * 2019-10-18 2022-06-10 지멘스 에너지 글로벌 게엠베하 운트 코. 카게 2개의 로터 디스크들 사이에 배열된 로터 구성 요소를 구비한 로터

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US20070231127A1 (en) * 2006-03-30 2007-10-04 Snecma Device for attaching ring sectors around a turbine rotor of a turbomachine
US7594792B2 (en) * 2005-04-27 2009-09-29 Snecma Sealing device for a chamber of a turbomachine, and aircraft engine equipped with said sealing device
US7686577B2 (en) * 2006-11-02 2010-03-30 Siemens Energy, Inc. Stacked laminate fiber wrapped segment

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FR951446A (fr) 1942-09-01 1949-10-25 Dynamit Nobel Ag Traverse de chemin de fer
FR2780443B1 (fr) * 1998-06-25 2000-08-04 Snecma Anneau de stator de turbine haute pression d'une turbomachine
JP4269829B2 (ja) * 2003-07-04 2009-05-27 株式会社Ｉｈｉ シュラウドセグメント
DE102005013798A1 (de) * 2005-03-24 2006-09-28 Alstom Technology Ltd. Wärmestausegment zum Abdichten eines Strömungskanals einer Strömungsrotationsmaschine
US7452183B2 (en) * 2005-08-06 2008-11-18 General Electric Company Thermally compliant turbine shroud assembly
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
FR2913717A1 (fr) * 2007-03-15 2008-09-19 Snecma Propulsion Solide Sa Ensemble d'anneau de turbine pour turbine a gaz

Patent Citations (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US7594792B2 (en) * 2005-04-27 2009-09-29 Snecma Sealing device for a chamber of a turbomachine, and aircraft engine equipped with said sealing device
US20070231127A1 (en) * 2006-03-30 2007-10-04 Snecma Device for attaching ring sectors around a turbine rotor of a turbomachine
US7686577B2 (en) * 2006-11-02 2010-03-30 Siemens Energy, Inc. Stacked laminate fiber wrapped segment

Cited By (13)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US9416676B2 (en) * 2013-09-06 2016-08-16 MTU Aero Engines AG Gas turbine
US20150192026A1 (en) * 2013-09-06 2015-07-09 MTU Aero Engines AG Gas turbine
US10100659B2 (en) 2014-12-16 2018-10-16 Rolls-Royce North American Technologies Inc. Hanger system for a turbine engine component
US10273825B2 (en) * 2015-05-15 2019-04-30 Rolls-Royce Plc Wall cooling arrangement for a gas turbine engine
US20160333784A1 (en) * 2015-05-15 2016-11-17 Rolls-Royce Plc A wall cooling arrangement for a gas turbine engine
US20180037511A1 (en) * 2015-06-10 2018-02-08 Ihi Corporation Turbine
US10597334B2 (en) * 2015-06-10 2020-03-24 Ihi Corporation Turbine comprising turbine stator vanes of a ceramic matrix composite attached to a turbine case
US20190040758A1 (en) * 2015-10-05 2019-02-07 Safran Aircraft Engines Turbine ring assembly with axial retention
US10787924B2 (en) * 2015-10-05 2020-09-29 Safran Aircraft Engines Turbine ring assembly with axial retention
US10378386B2 (en) * 2015-12-18 2019-08-13 Safran Aircraft Engines Turbine ring assembly with support when cold and when hot
US10605120B2 (en) * 2016-09-27 2020-03-31 Safran Aircraft Engines Turbine ring assembly that can be set while cold
US12037926B2 (en) 2020-06-18 2024-07-16 Siemens Energy Global GmbH & Co. KG Rotor comprising a rotor component arranged between two rotor discs
CN113898414A (zh) * 2021-12-09 2022-01-07 成都中科翼能科技有限公司 一种燃气轮机高压转子防热振动变形的补强结构

Also Published As

Publication number Publication date
CA2777370A1 (fr) 2011-06-23
RU2012130351A (ru) 2014-01-27
EP2513428A1 (de) 2012-10-24
WO2011073570A1 (fr) 2011-06-23
FR2954400A1 (fr) 2011-06-24
BR112012010257A2 (pt) 2016-03-29
CN102667066A (zh) 2012-09-12
FR2954400B1 (fr) 2012-03-09

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AS Assignment

Owner name: SNECMA, FRANCE

Free format text: ASSIGNMENT OF ASSIGNORS INTEREST;ASSIGNORS:BERCHE, EMMANUEL;PHILIPPOT, VINCENT;REEL/FRAME:028265/0287

Effective date: 20120313

STCB Information on status: application discontinuation

Free format text: ABANDONED -- FAILURE TO RESPOND TO AN OFFICE ACTION