US20120128482A1 - Outer shell sector for a bladed ring for an aircraft turbomachine stator, including vibration damping shims - Google Patents

Outer shell sector for a bladed ring for an aircraft turbomachine stator, including vibration damping shims Download PDF

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Publication number
US20120128482A1
US20120128482A1 US13/386,496 US201013386496A US2012128482A1 US 20120128482 A1 US20120128482 A1 US 20120128482A1 US 201013386496 A US201013386496 A US 201013386496A US 2012128482 A1 US2012128482 A1 US 2012128482A1
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US
United States
Prior art keywords
sector
outer shell
vibration damping
shim
elementary sectors
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
US13/386,496
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English (en)
Inventor
Laurent Gilles DEZOUCHE
Patrick Edmond Kapala
Samir Zaidi
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: DEZOUCHE, LAURENT GILLES, KAPALA, PATRICK EDMOND, ZAIDI, SAMIR
Publication of US20120128482A1 publication Critical patent/US20120128482A1/en
Abandoned legal-status Critical Current

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Classifications

    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F04—POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
    • F04D—NON-POSITIVE-DISPLACEMENT PUMPS
    • F04D29/00—Details, component parts, or accessories
    • F04D29/40—Casings; Connections of working fluid
    • F04D29/52—Casings; Connections of working fluid for axial pumps
    • F04D29/54—Fluid-guiding means, e.g. diffusers
    • F04D29/541—Specially adapted for elastic fluid pumps
    • F04D29/542—Bladed diffusers
    • 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
    • F01D11/001—Preventing or minimising internal leakage of working-fluid, e.g. between stages for sealing space between stator blade and rotor
    • 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
    • F01D25/04—Antivibration arrangements
    • F01D25/06—Antivibration arrangements for preventing blade vibration
    • 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
    • F04—POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
    • F04D—NON-POSITIVE-DISPLACEMENT PUMPS
    • F04D29/00—Details, component parts, or accessories
    • F04D29/60—Mounting; Assembling; Disassembling
    • F04D29/64—Mounting; Assembling; Disassembling of axial pumps
    • F04D29/644—Mounting; Assembling; Disassembling of axial pumps especially adapted for elastic fluid pumps
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F04—POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
    • F04D—NON-POSITIVE-DISPLACEMENT PUMPS
    • F04D29/00—Details, component parts, or accessories
    • F04D29/66—Combating cavitation, whirls, noise, vibration or the like; Balancing
    • F04D29/661—Combating cavitation, whirls, noise, vibration or the like; Balancing especially adapted for elastic fluid pumps
    • F04D29/668—Combating cavitation, whirls, noise, vibration or the like; Balancing especially adapted for elastic fluid pumps damping or preventing mechanical vibrations
    • 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
    • F01D9/042—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles forming ring or sector fixing blades to stators

Definitions

  • This invention generally relates to an aircraft turbomachine, preferably of the turbojet or turboprop type.
  • the invention relates to the compressor or turbine stator of such a turbomachine, and more precisely to a bladed ring sector comprising a plurality of stator blades and two concentric shells supporting the blades and designed to radially delimit a primary flow passing through the turbomachine, inwards and outwards respectively.
  • a bladed ring is usually made using several sectors arranged end to end, is usually used in the compressor or the turbine as a guide vane or a nozzle.
  • Turbomachines usually comprise a low pressure compressor, a high pressure compressor, a combustion chamber, a high pressure turbine and a low pressure turbine, in series. Compressors and turbines comprise several rows of mobile blades at a circumferential spacing, these rows being separated by rows of fixed blades also at a circumferential spacing.
  • high dynamic stresses are applied to the guide vanes and nozzles.
  • Technological progress leads to a reduction in the number of stages for equal or better performances, resulting in a higher load for each stage.
  • changes to production technologies have led to a reduction in the number of parts, which reduces the damping effect of connections between parts. This is the case particularly when an abradable cartridge brazing technology is used which eliminates a large source of dissipation of vibration energy.
  • Document FR-A-2 902 843 discloses a means of solving this vibration problem by breaking the outer shell sector down into elementary sectors at a fixed spacing from each other along the tangential direction by the use of slits or radial cuts, oblique or in another direction, each elementary sector supporting a single blade of the bladed ring sector. Furthermore, damping inserts in the form of strips are inserted between the elementary sectors.
  • the operating principle is based on the introduction of a stiffness non-linearity in the dynamic behaviour of the structure. This non-linearity is triggered by a threshold vibration level of the system. This vibration activity causes a relative movement between the elementary sectors of the blades and the damping inserts.
  • the purpose of the invention is to at least partially overcome the problems mentioned above that arise with embodiments according to prior art.
  • the first purpose of the invention to achieve this is an assembly forming an outer shell sector for a bladed ring sector that will be used on a compressor or turbine stator in an aircraft turbomachine, said outer shell sector comprising firstly a plurality of elementary sectors at a spacing from each other along a tangential direction of said assembly, and secondly vibration damping shims, each of them being inserted between two elementary sectors associated with it, placed directly consecutively along said tangential direction.
  • each vibration damping shim is approximately the same as the profile of the elementary sectors.
  • said shim bears in contact with two parallel plane friction surfaces facing each other along said tangential direction and provided on said two elementary sectors associated with said shim, and said shim has two complementary plane friction surfaces parallel to each other and cooperating with the two corresponding friction surfaces of the elementary sectors.
  • the plane contacts between the friction surfaces and the complementary friction surfaces give satisfactory damping of vibrations by friction. It is also possible to make the two friction surfaces simultaneously during a single machining operation, for example by a single cutting operation, in order to obtain straight slits, in other words slits in a determined plane, inside which the corresponding shims will subsequently be housed. This makes it very much simpler to fabricate the assembly according to the invention, which results in a significant cost and time saving.
  • said shim is provided with hooks to hold it in place on the compressor or turbine stator, therefore these hooks have the same profile as the hooks fixed on the elementary sectors.
  • the elementary sectors are separated from each other by radial slits completely filled in by said vibration damping shims.
  • said vibration damping shims extend approximately along an axial or oblique direction of said assembly.
  • FIG. 1 Another purpose of this invention applies to a bladed ring sector designed to be installed on a compressor or turbine stator of an aircraft turbomachine comprising an assembly forming an outer shell sector like that described above, an inner shell sector and a plurality of blades at a tangential spacing from each other and inserted between the assembly forming the outer shell sector and the inner shell sector.
  • each elementary sector will carry a single stator blade, or possibly several blades, without going outside the scope of the invention.
  • the bladed ring may form a guide vane of a compressor or a nozzle of a turbine.
  • the ring sector preferably extends around an angular range of between 5 and 60°, but can be as much as 360° so as to form the entire bladed ring.
  • Another purpose of the invention is an aircraft turbomachine comprising a compressor or turbine stator equipped with at least one bladed ring sector like that described above.
  • FIG. 1 shows a diagrammatic sectional view of a turbomachine that will be equipped with one or several bladed ring sectors according to this invention
  • FIG. 2 shows a sectional view representing part of the high pressure compressor of the turbomachine shown in FIG. 1 , and including a bladed ring sector according to this invention
  • FIG. 3 shows a perspective view of the bladed ring sector shown in the previous figure, the sector being in the form of a preferred embodiment of this invention
  • FIG. 4 shows an axial view of part of the bladed ring sector shown in the previous figure
  • FIG. 5 shows a profile view of the shims and the elementary sectors of the bladed ring sector shown in the previous figures, along line V-V in FIG. 4 ;
  • FIGS. 6 a to 6 c show views diagrammatically showing the different steps in a fabrication process of the bladed ring sector shown in the previous figures.
  • FIG. 1 shows an aircraft turbojet 100 to which the invention is applicable. It comprises, in order along the upstream to downstream direction, a low pressure compressor 2 , a high pressure compressor 4 , an annular combustion chamber 6 , a high pressure turbine 8 and a low pressure turbine 10 .
  • FIG. 2 shows part of the high pressure compressor 4 .
  • the compressor comprises rows 14 of stator blades and rows 16 of rotor blades alternating on an axial direction parallel to the axis 12 of the compressor.
  • the stator blades 18 distributed circumferentially/tangentially around the axis 12 , are included in a part of the stator called the bladed ring 20 , preferably constructed in sectors along the circumferential direction 22 .
  • a bladed ring sector 20 it being understood that this sector 20 preferably extends over an angular range of between 5 and 60°, but possibly as much as 360° so as to form the entire bladed ring.
  • the sector 20 therefore forming all or part of a turbine nozzle or a compressor guide vane, comprises an inner shell sector 24 forming the inner surface radially delimiting a primary annular flow 26 passing through the turbomachine, this shell sector 24 supporting the fixed roots of the stator blades 18 .
  • the sector 20 also comprises an assembly forming an outer shell sector 28 forming the outer surface radially delimiting the primary annular flow, and supporting the fixed heads of the blades 18 .
  • the sector 20 also comprises known additional elements fitted on the shell sector 24 , such as a radially internal abradable coating 29 forming the annular sealing track contacted by a sealing device 31 supported by the rotor stage 16 supporting the rotating blades and arranged on the downstream side of the sector 20 concerned.
  • the rotating sealing device 31 is a known labyrinth or lip seal type sealing device.
  • FIG. 3 shows the bladed ring sector 20 .
  • the entire turbine nozzle or compressor guide vane is obtained by end to end placement of a plurality of these sectors 20 , therefore each forming an angular or circumferential portion of this bladed ring.
  • the angular sectors 20 (only one of which can be seen in FIG. 3 ) are preferably deprived of any rigid direct mechanical links connecting them to each other, their adjacent ends being simply placed facing each other with or without clearance.
  • the figures show that the inner ring sector 24 is made in a single part and is not segmented.
  • the assembly 28 forming the outer shell sector 28 is segmented into elementary sectors 30 at a spacing from each other along the tangential direction 22 , by straight radial or slightly oblique slits 32 , therefore creating clearances between the directly consecutive sectors 30 .
  • Each slit 32 is made along a median straight line between two directly consecutive blades 18 , such that each elementary sector 30 supports a single fixed stator blade 18 .
  • One of the two elementary sectors 30 located at the ends of the sector 20 supports a rotation stop 33 projecting radially outwards and that will cooperate with another part of the compressor stator in a known manner.
  • the assembly 28 also comprises vibration damping shims 34 housed between directly consecutive elementary sectors 30 .
  • each vibration damping shim is housed between two plane parallel friction surfaces 38 facing each other along the tangential direction 22 , and provided on the corresponding tangential ends facing each other on the two elementary sectors associated with the shim.
  • each shim has two complementary plane friction surfaces 40 parallel to each other and also parallel and in contact with the two corresponding plane friction surfaces 38 with which they cooperate.
  • each shim 34 is squeezed between two directly consecutive elementary sectors 30 , having a shape complementary with the shape of the friction surfaces 38 .
  • the contact between the two friction surfaces 38 , 40 of each pair is preferably obtained as soon as the shim 34 is put into position between its two associated elementary sectors 30 .
  • the shims 34 thus apply forces oriented approximately along the tangential direction in contact with the friction surfaces 38 of the elementary sectors, with their complementary plane friction surfaces 40 . These forces are advantageously increased during operation by the additional application of the tangential component of aerodynamic forces applied on the stator blades, on the elementary sectors.
  • profile refers to the global shape of the element seen along the tangential direction 22 , although a sectional view is shown in FIG. 5 .
  • each shim 34 acts as the outer radial delimitation of the air flowpath. Consequently, the global annular delimitation surface of the air flowpath composed of the sequence of these surfaces 46 formed on the shims 34 and the sectors 30 , is approximately continuous from an aerodynamic point of view because there is no step between the successive surfaces 46 .
  • Each shim 34 and each sector 30 also comprises hooks to hold it in place on another part of the compressor stator, and more precisely a fixing hook 48 projecting forwards, and a fixing hook 50 projecting backwards. As shown in FIG. 2 , the hooks 48 , 50 are housed in the corresponding annular slits 52 , 54 provided in another part of the compressor stator, to fix the sector 20 onto this other part of the stator.
  • the shims 34 entirely filling in the slits 32 , perform a vibration damping function by friction in contact with the friction surfaces 38 , based on the physical principle described above for the shims disclosed in document FR-A-2 902 843. They also perform a seal function, and a function to allow the tangential component of aerodynamic forces applied on the stator blades to pass through. More generally in this respect, each shim 34 is capable of transmitting tangential forces between the two elementary sectors 30 between which it is inserted.
  • the natures of the materials used for the elementary sectors 30 and for the shims 34 are approximately the same, preferably metallic, and are chosen such that the shims wear preferentially rather than the elementary sectors 30 .
  • the ratio between the extent of each shim and the extent of each elementary sector along the tangential direction that also correspond to the thicknesses is between 0.5 and 1.
  • FIGS. 6 a to 6 c diagrammatically show a process for fabrication of the bladed ring sector 20 .
  • a single-piece assembly 100 is made by pouring or machining forming the inner shell sector 24 , the outer shell sector 28 and the stator blades 18 .
  • the next step is to make straight radial slits 32 on the outer shell sector 28 so as to obtain the elementary sectors 30 as shown diagrammatically in FIG. 6 b , by simple and inexpensive machining. For example, these slits 32 can be made simply by cutting the sector 28 .
  • FIG. 6 c shows the final step that consists of putting the vibration damping shims 34 into position in the slits 32 forming the friction surfaces, simply by sliding the shims into their corresponding holes.

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  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Structures Of Non-Positive Displacement Pumps (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
US13/386,496 2009-07-31 2010-07-29 Outer shell sector for a bladed ring for an aircraft turbomachine stator, including vibration damping shims Abandoned US20120128482A1 (en)

Applications Claiming Priority (3)

Application Number Priority Date Filing Date Title
FR0955439 2009-07-31
FR0955439A FR2948736B1 (fr) 2009-07-31 2009-07-31 Secteur de virole exterieure pour couronne aubagee de stator de turbomachine d'aeronef, comprenant des cales amortisseuses de vibrations
PCT/EP2010/061037 WO2011012679A2 (fr) 2009-07-31 2010-07-29 Secteur de virole exterieure pour couronne aubagee de stator de turbomachine d'aeronef, comprenant des cales amortisseuses de vibrations

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US20120128482A1 true US20120128482A1 (en) 2012-05-24

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US13/386,496 Abandoned US20120128482A1 (en) 2009-07-31 2010-07-29 Outer shell sector for a bladed ring for an aircraft turbomachine stator, including vibration damping shims

Country Status (9)

Country Link
US (1) US20120128482A1 (zh)
EP (1) EP2459884B1 (zh)
JP (1) JP5697667B2 (zh)
CN (1) CN102472297A (zh)
BR (1) BR112012002304A2 (zh)
CA (1) CA2769217A1 (zh)
FR (1) FR2948736B1 (zh)
RU (1) RU2537997C2 (zh)
WO (1) WO2011012679A2 (zh)

Cited By (8)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US20120201684A1 (en) * 2011-02-08 2012-08-09 Andy Turko Mate face brazing for turbine components
US20140001285A1 (en) * 2012-06-29 2014-01-02 General Electric Company Nozzle, a nozzle hanger, and a ceramic to metal attachment system
US9334756B2 (en) 2012-09-28 2016-05-10 United Technologies Corporation Liner and method of assembly
US20160230576A1 (en) * 2015-02-05 2016-08-11 Rolls-Royce North American Technologies, Inc. Vane assemblies for gas turbine engines
US10066548B2 (en) 2013-03-15 2018-09-04 United Technologies Corporation Acoustic liner with varied properties
US10337527B2 (en) 2014-11-28 2019-07-02 Safran Aircraft Engines Turbomachine blade, comprising intersecting partitions for circulation of air in the direction of the trailing edge
US20210156271A1 (en) * 2019-11-21 2021-05-27 United Technologies Corporation Vane with collar
FR3119196A1 (fr) * 2021-01-27 2022-07-29 Safran Aircraft Engines Rangée annulaire sectorisée d’aubes fixes

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Publication number Priority date Publication date Assignee Title
FR2971022B1 (fr) 2011-02-02 2013-01-04 Snecma Etage redresseur de compresseur pour une turbomachine
DE102013212252A1 (de) * 2013-06-26 2014-12-31 Siemens Aktiengesellschaft Turbine und Verfahren zur Anstreiferkennung
FR3008455B1 (fr) * 2013-07-09 2015-08-21 Snecma Redresseur pour compresseur comportant des moyens de rattrapage de jeu
CN104440153B (zh) * 2014-11-04 2017-06-06 中国南方航空工业(集团)有限公司 机匣内叶片加工阻尼装置
JP6689117B2 (ja) * 2016-03-31 2020-04-28 三菱日立パワーシステムズ株式会社 軸流回転機械に装備される静翼環及び軸流回転機械
CN106988794B (zh) * 2017-06-02 2018-12-14 中国航发南方工业有限公司 静子组合件夹紧方法及静子组合件
CN107747563B (zh) * 2017-09-30 2020-04-10 中国航发沈阳发动机研究所 带阻尼的风扇机匣
FR3115819B1 (fr) * 2020-11-02 2023-04-14 Safran Aircraft Engines Ensemble de stator de turbomachine d’aéronef, comprenant une structure externe formée de deux tronçons annulaires entourant une couronne aubagée de stator

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US20070297900A1 (en) * 2006-06-23 2007-12-27 Snecma Sector of a compressor guide vanes assembly or a sector of a turbomachine nozzle assembly
US20080206063A1 (en) * 2007-02-27 2008-08-28 Lynn Charles Gagne Method and apparatus for assembling blade shims

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FR2831615B1 (fr) * 2001-10-31 2004-01-02 Snecma Moteurs Redresseur fixe sectorise pour compresseur d'une turbomachine
US6733237B2 (en) * 2002-04-02 2004-05-11 Watson Cogeneration Company Method and apparatus for mounting stator blades in axial flow compressors
EP1510654A1 (de) * 2003-08-25 2005-03-02 Siemens Aktiengesellschaft Einstückige Turbinenbeschaufelungseinheit und Verfahren zur Herstellung einer einstückigen Turbinenbeschaufelungseinheit
US7104752B2 (en) * 2004-10-28 2006-09-12 Florida Turbine Technologies, Inc. Braided wire damper for segmented stator/rotor and method
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US6984108B2 (en) * 2002-02-22 2006-01-10 Drs Power Technology Inc. Compressor stator vane
US20070297900A1 (en) * 2006-06-23 2007-12-27 Snecma Sector of a compressor guide vanes assembly or a sector of a turbomachine nozzle assembly
US20080206063A1 (en) * 2007-02-27 2008-08-28 Lynn Charles Gagne Method and apparatus for assembling blade shims

Cited By (13)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US20120201684A1 (en) * 2011-02-08 2012-08-09 Andy Turko Mate face brazing for turbine components
US9610644B2 (en) * 2011-02-08 2017-04-04 United Technologies Corporation Mate face brazing for turbine components
US20140001285A1 (en) * 2012-06-29 2014-01-02 General Electric Company Nozzle, a nozzle hanger, and a ceramic to metal attachment system
US9546557B2 (en) * 2012-06-29 2017-01-17 General Electric Company Nozzle, a nozzle hanger, and a ceramic to metal attachment system
US9334756B2 (en) 2012-09-28 2016-05-10 United Technologies Corporation Liner and method of assembly
USRE48980E1 (en) 2013-03-15 2022-03-22 Raytheon Technologies Corporation Acoustic liner with varied properties
US10066548B2 (en) 2013-03-15 2018-09-04 United Technologies Corporation Acoustic liner with varied properties
US10337527B2 (en) 2014-11-28 2019-07-02 Safran Aircraft Engines Turbomachine blade, comprising intersecting partitions for circulation of air in the direction of the trailing edge
US20160230576A1 (en) * 2015-02-05 2016-08-11 Rolls-Royce North American Technologies, Inc. Vane assemblies for gas turbine engines
US10655482B2 (en) * 2015-02-05 2020-05-19 Rolls-Royce Corporation Vane assemblies for gas turbine engines
US20210156271A1 (en) * 2019-11-21 2021-05-27 United Technologies Corporation Vane with collar
US11242762B2 (en) * 2019-11-21 2022-02-08 Raytheon Technologies Corporation Vane with collar
FR3119196A1 (fr) * 2021-01-27 2022-07-29 Safran Aircraft Engines Rangée annulaire sectorisée d’aubes fixes

Also Published As

Publication number Publication date
CN102472297A (zh) 2012-05-23
EP2459884A2 (fr) 2012-06-06
EP2459884B1 (fr) 2018-06-27
FR2948736B1 (fr) 2011-09-23
WO2011012679A3 (fr) 2011-04-21
JP5697667B2 (ja) 2015-04-08
RU2537997C2 (ru) 2015-01-10
RU2012107522A (ru) 2013-09-10
BR112012002304A2 (pt) 2016-05-31
JP2013501181A (ja) 2013-01-10
CA2769217A1 (en) 2011-02-03
WO2011012679A2 (fr) 2011-02-03
FR2948736A1 (fr) 2011-02-04

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AS Assignment

Owner name: SNECMA, FRANCE

Free format text: ASSIGNMENT OF ASSIGNORS INTEREST;ASSIGNORS:DEZOUCHE, LAURENT GILLES;KAPALA, PATRICK EDMOND;ZAIDI, SAMIR;REEL/FRAME:027591/0241

Effective date: 20120112

STCB Information on status: application discontinuation

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