EP1265035B1 - Doppelbefestigung einer Turbinenbrennkammer aus keramischem Matrix-Verbundwerkstoff - Google Patents

Doppelbefestigung einer Turbinenbrennkammer aus keramischem Matrix-Verbundwerkstoff Download PDF

Info

Publication number
EP1265035B1
EP1265035B1 EP02291364A EP02291364A EP1265035B1 EP 1265035 B1 EP1265035 B1 EP 1265035B1 EP 02291364 A EP02291364 A EP 02291364A EP 02291364 A EP02291364 A EP 02291364A EP 1265035 B1 EP1265035 B1 EP 1265035B1
Authority
EP
European Patent Office
Prior art keywords
metal
combustion chamber
turbomachine according
tabs
composite material
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.)
Expired - Lifetime
Application number
EP02291364A
Other languages
English (en)
French (fr)
Other versions
EP1265035A1 (de
Inventor
Didier Hernandez
Gwénaelle Calvez
Alexandre Forestier
Eric Conete
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 EP1265035A1 publication Critical patent/EP1265035A1/de
Application granted granted Critical
Publication of EP1265035B1 publication Critical patent/EP1265035B1/de
Anticipated expiration legal-status Critical
Expired - Lifetime legal-status Critical Current

Links

Images

Classifications

    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
    • F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
    • F23R3/42—Continuous combustion chambers using liquid or gaseous fuel characterised by the arrangement or form of the flame tubes or combustion chambers
    • F23R3/60—Support structures; Attaching or mounting means
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
    • F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
    • F23R3/007—Continuous combustion chambers using liquid or gaseous fuel constructed mainly of ceramic components
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05B—INDEXING SCHEME RELATING TO WIND, SPRING, WEIGHT, INERTIA OR LIKE MOTORS, TO MACHINES OR ENGINES FOR LIQUIDS COVERED BY SUBCLASSES F03B, F03D AND F03G
    • F05B2230/00—Manufacture
    • F05B2230/60—Assembly methods
    • F05B2230/604—Assembly methods using positioning or alignment devices for aligning or centering, e.g. pins
    • F05B2230/606—Assembly methods using positioning or alignment devices for aligning or centering, e.g. pins using maintaining alignment while permitting differential dilatation

Definitions

  • the present invention relates to the specific field of turbomachines and is more particularly concerned with the problem of mounting a combustion chamber made of a composite material of the CMC (ceramic matrix composite) type in the metal casing of a turbomachine.
  • CMC ceramic matrix composite
  • the high pressure turbine including its inlet nozzle (HPT nozzle), the combustion chamber and the casing (or envelope) of this chamber are made of the same material, generally of the type metallic.
  • HPT nozzle inlet nozzle
  • the combustion chamber and the casing (or envelope) of this chamber are made of the same material, generally of the type metallic.
  • the use of a metal chamber is from a thermal point of view totally inadequate and it must be resorted to a chamber based on CMC type high temperature composite materials.
  • the difficulties of implementation and the cost of these materials mean that their use is most often limited to the combustion chamber itself, the inlet valve of the high pressure turbine and the casing then remaining more conventionally achieved.
  • metallic materials metal materials and composite materials have very different coefficients of thermal expansion. This results in particularly acute problems of connection with the housing and the combustion chamber and sealing at the distributor at the inlet of the high pressure turbine.
  • the US Patent 6,131,384 shows a turbomachine according to this prior art.
  • the present invention overcomes these disadvantages by proposing a mounting of the combustion chamber in the casing having the capacity to absorb the displacements induced by the differences in the expansion coefficients of these parts.
  • a turbomachine comprising, in an envelope of metallic material and in a direction F of gas flow, a fuel injection system, a composite material combustion chamber having a longitudinal axis, and a metal material distributor forming the fixed blade inlet stage of a high pressure turbine, characterized in that said composite material combustion chamber is held in position in said metal casing by a plurality of flexible metal tabs having first and second ends, said first ends being interconnected by a metal flange crown fixed to said metal shell by first fixing means and said second ends being each fixed jointly by second fixing means on the one hand to said chamber of composite material combustion and secondly at one end of a composite material wall whose other end forms a support plane for a sealing element integral with said distributor and ensuring the sealing of the gas stream between said combustion chamber and said distributor, the flexibility of said fixing lugs permitting at high temperatures radial free expansion of said combustion chamber made of composite material with respect to said metal casing.
  • the first and second fastening means are preferably constituted by a plurality of bolts.
  • the second fixing means can also be constituted by crimping elements.
  • said sealing element is of the type "circular seal”. It may comprise a plurality of calibrated leakage orifices.
  • said metal ring interconnecting said first ends of said flexible metal tabs is mounted between connecting flanges of these two parts.
  • said metal ring may be fixed directly to said annular casing by conventional fastening means.
  • said first ends of the fixing lugs may either be fixed by brazing (or welding) to said metal flange ring or form a single piece with this metal ring.
  • the distributor is fixed on a downstream portion 14b of the inner annular envelope of the turbomachine by first removable fastening means preferably constituted by a plurality of bolts 50 while resting on support means 49 integral with the outer annular envelope of the turbomachine.
  • Through-holes 54, 56 formed in the outer metal 46 and inner 48 metal platforms of the distributor 42 are furthermore provided for cooling the vanes 44 of the distributor at the inlet of the rotor of the high-pressure turbine from the oxidant compressed available at the outlet of the diffusion duct 18 and flowing in two flows F1, F2 on either side of the combustion chamber 24.
  • These attachment tabs are mounted for a first part of them (see the tab referenced 58) between the outer annular casing 12a, 12b and the outer axial wall 26 of the combustion chamber and for a second part (such as the tab 60) between the inner annular envelope 14a, 14b and the inner axial wall 28 of the combustion chamber.
  • the number of legs may, for example, be in number equal to that of the injection nozzles or equal to a multiple of this number.
  • Each flexible fastening clip of metal material which can have a substantially triangular shape as illustrated Figure 1A , or consist of a single blade (not shown constant width or not), is welded or soldered by a first end 62; 64 to a metal ring 66a, 66b forming a flange and secured integrally by first fixing means 52; 68 to one or the other (depending on its location) of the outer or inner metal annular envelopes.
  • This attachment flange is intended to facilitate the maintenance of these legs on the metal shells.
  • these tabs and the metal ring together form a single piece of metal in one piece.
  • this tab is fixed jointly by second fixing means 74, 76 on the one hand at a downstream end 88; 90 of the external axial walls 26 and inner 28 of the combustion chamber made of ceramic composite material and secondly at one end of a ceramic composite wall 78a; 78b disposed in the extension of each of the outer and inner axial walls, forming a kind of a second chamber portion, and the other end forms a support plane for a sealing element integral with the distributor and sealing the gas stream between the combustion chamber 24 and the distributor 42.
  • connection of the second ends of the tabs 70, 72 with the downstream ends of the walls of the combustion chamber and the first ends of the ceramic composite walls forming the second chamber portion is carried out by a simple bolting, preferably of prison nut type to facilitate a possible assembly / disassembly and correlatively limit the dimensions of the legs.
  • the metal ring 66a, 66b interconnecting the first ends 62, 64 of the tabs is in turn preferably taken between existing connecting flanges between the upstream portions 12a, 14a and downstream 12b, 14b of the inner and outer annular envelopes and maintained fixedly by the first attachment means 52, 68 which are preferably also bolt type.
  • washers ceramic composite material 74a; 76a to allow to "drown" the conical heads of the screws of the bolts forming the second attachment means 74; 76.
  • the tightness of the gas stream between the combustion chamber 24 and the distributor 42 is provided by a "flap" circular seal 80, 82 mounted in a groove 84, 86 of each of the outer 46 and inner 48 platforms. dispenser and which bears directly on the second end portion of the ceramic composite wall 78a; 78b forming a support plane for this seal circular sealing.
  • the seal is held in abutment against this second end of the composite wall by means of a resilient element, leaf spring type 92, 94, fixed on the distributor.
  • an omega-type circular seal 96 mounted in a circular groove 98 of a flange of the inner annular casing 14 in direct contact with the inner circular platform 48 of the distributor and secondly by another "flap" circular seal 100 mounted in a circular groove 102 of the outer circular platform of the distributor 46 and one end is in direct contact with a circular spoiler 104 of the downstream portion 12b of the outer annular envelope.
  • Figure 1B illustrates a first variant of the previous embodiment in which the attachment of the tabs (only the case of the tab 60 is illustrated) at the downstream end 90 of the combustion chamber 24 is performed by a crimped connection, the bolts 76 being replaced by crimping elements 76b.
  • the cooling being able to take place through the crimping elements, it is no longer necessary to provide calibrated orifices at the lamellae joints 80, 82.
  • the metal ring 66a forming a flange interconnecting by brazing (or welding) the first ends 62 of the fastening tabs 58 of the outer axial wall of the combustion chamber 26 is no longer mounted between flanges but itself brazed (or welded) ) at a polarizer 106 centered and resting on the outer annular casing 12.
  • the metal ring 66b forming a flange interconnecting by brazing (or welding) the first ends 64 of the fixing lugs 60 of the inner axial wall of the combustion chamber 28 is no longer mounted between flanges but simply fixed directly to the inner annular casing 14 by conventional fastening means 108, for example of the bolt type.
  • the flexibility of the fixing lugs makes it possible to withstand the thermal expansion gap occurring at the temperatures elevated between the composite material combustion chamber and the metal annular envelopes while ensuring the maintenance and positioning of the chamber.

Landscapes

  • Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Ceramic Engineering (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
  • Chimneys And Flues (AREA)
  • Gasket Seals (AREA)

Claims (9)

  1. Turbomaschine, die - in einer Hülle aus metallischem Werkstoff (12a, 12b; 14a, 14b) und entlang einer Strömungsrichtung F der Gase - eine Einheit zum Einspritzen eines Kraftstoffs (20, 22), eine Brennkammer aus Verbundwerkstoff (24) mit einer Längsachse (10) sowie einen Leitapparat aus metallischem Werkstoff (42), der die Eintrittsstufe mit Leitschaufeln (44) einer Hochdruckturbine bildet, umfaßt, dadurch gekennzeichnet daß die Brennkammer aus Verbundwerkstoff durch eine Vielzahl von flexiblen Metallaschen (58, 60), die erste Enden (62, 64) und zweite Enden (70, 72) aufweisen, in der Metallhülle in Position gehalten wird, wobei die genannten ersten Enden durch einen einen Flansch bildenden Metallkranz (66a, 66b) untereinander verbunden sind, der an der Metallhülle durch erste Befestigungsmittel (52; 68, 108) befestigt ist, und wobei die genannten zweiten Enden jeweils gemeinsam durch zweite Befestigungsmittel (74, 76) einerseits an der Brennkammer aus Verbundwerkstoff (26, 28) und andererseits an einem Ende einer Wand aus Verbundwerkstoff (78a, 78b) befestigt sind, deren anderes Ende eine Stützebene für ein Dichtungselement (80, 82) bildet, das mit dem Leitapparat fest verbunden ist und das die Dichtigkeit des Gaskanals zwischen der Brennkammer und dem Leitapparat sicherstellt, wobei die Flexibilität der genannten Metallbefestigungslaschen bei hohen Temperaturen eine freie radiale Ausdehnung der Brennkammer aus Verbundwerkstoff gegenüber der Metallhülle ermöglicht.
  2. Turbomaschine nach Anspruch 1, dadurch gekennzeichnet, daß die ersten und zweiten Befestigungsmittel von einer Vielzahl von Bolzen gebildet sind.
  3. Turbomaschine nach Anspruch 1, dadurch gekennzeichnet, daß - da die Metallhülle aus zwei Teilen (12a, 12b; 14a, 14b) gebildet ist - der Metallkranz, welcher die ersten Enden der flexiblen Metallaschen untereinander verbindet, zwischen Verbindungsflanschen dieser zwei Teile angebracht ist.
  4. Turbomaschine nach Anspruch 1, dadurch gekennzeichnet, daß der Metallkranz, welcher die ersten Enden der flexiblen Metallaschen untereinander verbindet, durch herkömmliche Befestigungsmittel (108) direkt an der ringförmigen Hülle befestigt ist.
  5. Turbomaschine nach Anspruch 1, dadurch gekennzeichnet, daß die ersten Enden der flexiblen Metallaschen an dem einen Flansch bildenden Metallkranz festgelötet (oder festgeschweißt) sind.
  6. Turbomaschine nach Anspruch 1, dadurch gekennzeichnet, daß die ersten Enden der flexiblen Metallaschen mit dem einen Flansch bildenden Metallkranz ein einziges Teil bilden.
  7. Turbomaschine nach Anspruch 1, dadurch gekennzeichnet, daß die zweiten Befestigungsmittel von Crimpelementen (76b) gebildet sind.
  8. Turbomaschine nach Anspruch 1, dadurch gekennzeichnet, daß das Dichtungselement vom Typ "Lamellen"-Runddichtung (80, 82) ist.
  9. Turbomaschine nach Anspruch 8, dadurch gekennzeichnet, daß die "Lamellen"-Dichtung eine Vielzahl von kalibrierten Entweichungsöffnungen (110) aufweist.
EP02291364A 2001-06-06 2002-06-04 Doppelbefestigung einer Turbinenbrennkammer aus keramischem Matrix-Verbundwerkstoff Expired - Lifetime EP1265035B1 (de)

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
FR0107372 2001-06-06
FR0107372A FR2825785B1 (fr) 2001-06-06 2001-06-06 Liaison de chambre de combustion cmc de turbomachine en deux parties

Publications (2)

Publication Number Publication Date
EP1265035A1 EP1265035A1 (de) 2002-12-11
EP1265035B1 true EP1265035B1 (de) 2008-02-13

Family

ID=8863994

Family Applications (1)

Application Number Title Priority Date Filing Date
EP02291364A Expired - Lifetime EP1265035B1 (de) 2001-06-06 2002-06-04 Doppelbefestigung einer Turbinenbrennkammer aus keramischem Matrix-Verbundwerkstoff

Country Status (5)

Country Link
US (1) US6675585B2 (de)
EP (1) EP1265035B1 (de)
JP (1) JP4097994B2 (de)
DE (1) DE60224956T2 (de)
FR (1) FR2825785B1 (de)

Cited By (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN110822482A (zh) * 2019-11-28 2020-02-21 中国航发沈阳黎明航空发动机有限责任公司 一种中低热值气体和液体双燃料喷嘴及燃料切换方法

Families Citing this family (29)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP1381811A1 (de) * 2001-04-27 2004-01-21 Siemens Aktiengesellschaft Brennkammer, insbesondere einer gasturbine
FR2840974B1 (fr) * 2002-06-13 2005-12-30 Snecma Propulsion Solide Anneau d'etancheite pour cahmbre de combustion et chambre de combustion comportant un tel anneau
US6895761B2 (en) * 2002-12-20 2005-05-24 General Electric Company Mounting assembly for the aft end of a ceramic matrix composite liner in a gas turbine engine combustor
FR2855249B1 (fr) * 2003-05-20 2005-07-08 Snecma Moteurs Chambre de combustion ayant une liaison souple entre un fond de chambre et une paroi de chambre
FR2860039B1 (fr) * 2003-09-19 2005-11-25 Snecma Moteurs Realisation de l'etancheite dans un turboreacteur pour le prelevement cabine par joints double sens a lamelles
FR2871847B1 (fr) * 2004-06-17 2006-09-29 Snecma Moteurs Sa Montage d'un distributeur de turbine sur une chambre de combustion a parois en cmc dans une turbine a gaz
FR2871845B1 (fr) * 2004-06-17 2009-06-26 Snecma Moteurs Sa Montage de chambre de combustion de turbine a gaz avec distributeur integre de turbine haute pression
FR2871846B1 (fr) * 2004-06-17 2006-09-29 Snecma Moteurs Sa Chambre de combustion en cmc de turbine a gaz supportee dans un carter metallique par des organes de liaison en cmc
US7421842B2 (en) * 2005-07-18 2008-09-09 Siemens Power Generation, Inc. Turbine spring clip seal
FR2892181B1 (fr) * 2005-10-18 2008-02-01 Snecma Sa Fixation d'une chambre de combustion a l'interieur de son carter
US7775050B2 (en) * 2006-10-31 2010-08-17 General Electric Company Method and apparatus for reducing stresses induced to combustor assemblies
FR2930628B1 (fr) * 2008-04-24 2010-04-30 Snecma Chambre annulaire de combustion pour turbomachine
FR2935753B1 (fr) * 2008-09-08 2011-07-01 Snecma Propulsion Solide Liaisons souples a butee pour fixation de piece en cmc
US9234431B2 (en) * 2010-07-20 2016-01-12 Siemens Energy, Inc. Seal assembly for controlling fluid flow
US8322141B2 (en) * 2011-01-14 2012-12-04 General Electric Company Power generation system including afirst turbine stage structurally incorporating a combustor
US9335051B2 (en) * 2011-07-13 2016-05-10 United Technologies Corporation Ceramic matrix composite combustor vane ring assembly
FR2989426B1 (fr) * 2012-04-11 2014-03-28 Snecma Turbomachine, telle qu'un turboreacteur ou un turbopropulseur d'avion
FR2992687B1 (fr) * 2012-06-28 2014-07-18 Snecma Moteur a turbine a gaz comprenant une piece composite et une piece metallique reliees par un dispositif de fixation souple
CA2922569C (en) 2013-09-11 2018-02-20 General Electric Company Spring loaded and sealed ceramic matrix composite combustor liner
WO2015038293A1 (en) 2013-09-11 2015-03-19 United Technologies Corporation Combustor liner
US10378771B2 (en) 2016-02-25 2019-08-13 General Electric Company Combustor assembly
US10281153B2 (en) 2016-02-25 2019-05-07 General Electric Company Combustor assembly
US10519811B2 (en) * 2016-10-04 2019-12-31 United Technologies Corporation Flange heat shield
US10550725B2 (en) * 2016-10-19 2020-02-04 United Technologies Corporation Engine cases and associated flange
ES2760550T3 (es) * 2017-04-07 2020-05-14 MTU Aero Engines AG Disposición de junta para una turbina de gas
FR3084731B1 (fr) * 2019-02-19 2020-07-03 Safran Aircraft Engines Chambre de combustion pour une turbomachine
CN114413285B (zh) * 2022-01-29 2023-03-21 中国航发湖南动力机械研究所 一种大弯管密封结构
CN115405370B (zh) * 2022-11-03 2023-03-10 中国航发沈阳发动机研究所 一种半弹性涡轮外环结构
CN115717568B (zh) * 2022-12-27 2024-08-30 西安鑫垚陶瓷复合材料股份有限公司 一种陶瓷基复合材料混合器的安装装置

Family Cites Families (14)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2509503A (en) * 1946-02-12 1950-05-30 Lucas Ltd Joseph Combustion chamber for prime movers
US4030875A (en) * 1975-12-22 1977-06-21 General Electric Company Integrated ceramic-metal combustor
GB1570875A (en) * 1977-03-16 1980-07-09 Lucas Industries Ltd Combustion equipment
CH633351A5 (de) * 1978-11-09 1982-11-30 Sulzer Ag Waermedehnungen nachgebende abdichtung einer ringbrennkammer fuer eine gasturbine.
US5181377A (en) * 1991-04-16 1993-01-26 General Electric Company Damped combustor cowl structure
CA2070511C (en) * 1991-07-22 2001-08-21 Steven Milo Toborg Turbine nozzle support
FR2686683B1 (fr) * 1992-01-28 1994-04-01 Snecma Turbomachine a chambre de combustion demontable.
US5291733A (en) * 1993-02-08 1994-03-08 General Electric Company Liner mounting assembly
US5564271A (en) * 1994-06-24 1996-10-15 United Technologies Corporation Pressure vessel fuel nozzle support for an industrial gas turbine engine
DE19745683A1 (de) * 1997-10-16 1999-04-22 Bmw Rolls Royce Gmbh Aufhängung einer ringförmigen Gasturbinen-Brennkammer
JP4031590B2 (ja) * 1999-03-08 2008-01-09 三菱重工業株式会社 燃焼器の尾筒シール構造及びその構造を用いたガスタービン
US6397603B1 (en) * 2000-05-05 2002-06-04 The United States Of America As Represented By The Secretary Of The Air Force Conbustor having a ceramic matrix composite liner
US6334298B1 (en) * 2000-07-14 2002-01-01 General Electric Company Gas turbine combustor having dome-to-liner joint
US6497104B1 (en) * 2000-10-30 2002-12-24 General Electric Company Damped combustion cowl structure

Cited By (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN110822482A (zh) * 2019-11-28 2020-02-21 中国航发沈阳黎明航空发动机有限责任公司 一种中低热值气体和液体双燃料喷嘴及燃料切换方法

Also Published As

Publication number Publication date
US20020184888A1 (en) 2002-12-12
DE60224956D1 (de) 2008-03-27
JP4097994B2 (ja) 2008-06-11
DE60224956T2 (de) 2009-02-05
FR2825785B1 (fr) 2004-08-27
FR2825785A1 (fr) 2002-12-13
US6675585B2 (en) 2004-01-13
EP1265035A1 (de) 2002-12-11
JP2003035418A (ja) 2003-02-07

Similar Documents

Publication Publication Date Title
EP1265035B1 (de) Doppelbefestigung einer Turbinenbrennkammer aus keramischem Matrix-Verbundwerkstoff
EP1265034B1 (de) Befestigung einer Turbinenbrennkammer aus keramischem Matrix-Verbundswerkstoff mit gelöteten Befestigungsfüssen
EP1265037B1 (de) Verdünnungsluftlöcher verwendende Befestigung einer Turbinenbrennkammer aus keramischem Matrix-Verbundwerkstoff
EP1265036B1 (de) Elastische Befestigung einer Turbinenbrennkammer aus keramischem Matrix-Verbundwerkstoff in einem metallischen Gehäuse
EP1265030B1 (de) Befestigung einer Turbinenbrennkammer aus keramischem Matrix-Verbundwerkstoff mit flexiblen Verbindungsringen
EP1265032B1 (de) Gasturbinenbrennkammer aus Verbundwerkstoff mit keramischer Matrix
EP1265031B1 (de) Befestigung von metallischen Aufsätzen auf CMC-Turbomaschinenbrennkammerwänden
EP1265033B1 (de) Brennkammer mit einer Befestigungsanordnung der Kammerendwand
EP1734305B1 (de) Ringförmige Brennkammer für eine Turbine
EP1818615B1 (de) Ringförmige Brennkammer eines Turbotriebwerks
EP1607582B1 (de) Aufhängung einer Gasturbinenbrennkammer mit integriertem Turbinenleitapparat
EP2334909B1 (de) Dichtung zwischen einer brennkammer und einem turbinenleitapparat in einem turbomotor
EP1705340B1 (de) Verbindungsstück zwischen einem Kühlluftplenum und einer Statorschaufel in einer Turbomaschine
FR3004518A1 (fr) Chambre de combustion annulaire d'une turbomachine
EP1777460A1 (de) Befestigung einer Brennkammer im Inneren ihres Gehäuses
FR2825778A1 (fr) Liaison coulissante entre un systeme d'injection d'une chambre de combustion et un fond de cette chambre de combustion
FR2825782A1 (fr) Montage flottant radial de chambre de combustion cmc de turbomachine dans un carter metallique
FR3017928B1 (fr) Turbomachine a bride externe de chambre de combustion de type "sandwich"
FR3010774A1 (fr) Turbomachine a chambre de combustion maintenue par une couronne de fixation metallique
FR3111964A1 (fr) Assemblage d’une pièce de chambre de combustion par recouvrement par une autre pièce

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

17P Request for examination filed

Effective date: 20020610

AK Designated contracting states

Kind code of ref document: A1

Designated state(s): AT BE CH CY DE DK ES FI FR GB GR IE IT LI LU MC NL PT SE TR

AX Request for extension of the european patent

Free format text: AL;LT;LV;MK;RO;SI

AKX Designation fees paid

Designated state(s): DE FR GB

RAP1 Party data changed (applicant data changed or rights of an application transferred)

Owner name: SNECMA

GRAP Despatch of communication of intention to grant a patent

Free format text: ORIGINAL CODE: EPIDOSNIGR1

RIN1 Information on inventor provided before grant (corrected)

Inventor name: FORESTIER, ALEXANDRE

Inventor name: CONETE, ERIC

Inventor name: HERNANDEZ, DIDIER

Inventor name: CALVEZ, GWENAELLE

GRAS Grant fee paid

Free format text: ORIGINAL CODE: EPIDOSNIGR3

GRAA (expected) grant

Free format text: ORIGINAL CODE: 0009210

AK Designated contracting states

Kind code of ref document: B1

Designated state(s): DE FR GB

REG Reference to a national code

Ref country code: GB

Ref legal event code: FG4D

Free format text: NOT ENGLISH

REF Corresponds to:

Ref document number: 60224956

Country of ref document: DE

Date of ref document: 20080327

Kind code of ref document: P

GBT Gb: translation of ep patent filed (gb section 77(6)(a)/1977)

Effective date: 20080425

PLBE No opposition filed within time limit

Free format text: ORIGINAL CODE: 0009261

STAA Information on the status of an ep patent application or granted ep patent

Free format text: STATUS: NO OPPOSITION FILED WITHIN TIME LIMIT

26N No opposition filed

Effective date: 20081114

REG Reference to a national code

Ref country code: FR

Ref legal event code: PLFP

Year of fee payment: 15

REG Reference to a national code

Ref country code: FR

Ref legal event code: PLFP

Year of fee payment: 16

REG Reference to a national code

Ref country code: FR

Ref legal event code: CD

Owner name: SAFRAN AIRCRAFT ENGINES

Effective date: 20170719

REG Reference to a national code

Ref country code: FR

Ref legal event code: PLFP

Year of fee payment: 17

PGFP Annual fee paid to national office [announced via postgrant information from national office to epo]

Ref country code: FR

Payment date: 20210519

Year of fee payment: 20

Ref country code: DE

Payment date: 20210519

Year of fee payment: 20

PGFP Annual fee paid to national office [announced via postgrant information from national office to epo]

Ref country code: GB

Payment date: 20210519

Year of fee payment: 20

REG Reference to a national code

Ref country code: DE

Ref legal event code: R071

Ref document number: 60224956

Country of ref document: DE

REG Reference to a national code

Ref country code: GB

Ref legal event code: PE20

Expiry date: 20220603

PG25 Lapsed in a contracting state [announced via postgrant information from national office to epo]

Ref country code: GB

Free format text: LAPSE BECAUSE OF EXPIRATION OF PROTECTION

Effective date: 20220603