EP1265036B1 - Elastische Befestigung einer Turbinenbrennkammer aus keramischem Matrix-Verbundwerkstoff in einem metallischen Gehäuse - Google Patents

Elastische Befestigung einer Turbinenbrennkammer aus keramischem Matrix-Verbundwerkstoff in einem metallischen Gehäuse Download PDF

Info

Publication number
EP1265036B1
EP1265036B1 EP02291365A EP02291365A EP1265036B1 EP 1265036 B1 EP1265036 B1 EP 1265036B1 EP 02291365 A EP02291365 A EP 02291365A EP 02291365 A EP02291365 A EP 02291365A EP 1265036 B1 EP1265036 B1 EP 1265036B1
Authority
EP
European Patent Office
Prior art keywords
combustion chamber
turbomachine according
gasket
composite material
turbomachine
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
EP02291365A
Other languages
English (en)
French (fr)
Other versions
EP1265036A1 (de
Inventor
Pierre Camy
Benoît Carrere
Eric Conete
Alexandre Forestier
Georges Habarou
Didier Hernandez
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 EP1265036A1 publication Critical patent/EP1265036A1/de
Application granted granted Critical
Publication of EP1265036B1 publication Critical patent/EP1265036B1/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
  • Such a turbomachine is known from GB 1,570,875 .
  • the high pressure turbine including its inlet nozzle (HPT nozzle), the combustion chamber and the housing (also called envelope) of this chamber are made of the same material, generally metallic .
  • HPT nozzle inlet nozzle
  • the combustion chamber and the housing (also called envelope) of this chamber are made of the same material, generally 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 of these materials and their cost make 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 typically made in metallic materials.
  • metal materials and composite materials have very different coefficients of thermal expansion. This results in particularly acute problems of connection between the housing and the combustion chamber and interface at the distributor at the inlet of the high pressure turbine.
  • 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.
  • An object of the invention is also to provide a mounting which allows a simplification of the manufacture of the combustion chamber.
  • a turbomachine comprising, in an annular envelope of metallic material and in a direction F of flow of gas, a fuel injection assembly, an annular combustion chamber of composite material having a longitudinal axis and an annular distributor of metallic material forming the fixed blade inlet stage of a high pressure turbine, said chamber combustion device made of composite material being held in position in said annular metal casing by a plurality of tongues of elastic material, regularly distributed around said combustion chamber, characterized in that each of these tabs comprises three branches connected in a star, the ends of two of these three branches being fixed integrally to a downstream end of said combustion chamber of composite material, opposite to said injection system, respectively first and second fixing means and the end of the last of these three branches being fixed solidarily to said annular metallic envelope by third securing means, the elasticity of said fastening tabs allowing high temperatures free radial expansion of said combustion chamber relative to said annular envelope.
  • the first, second and third fixing means are each preferably constituted by a plurality of bolts.
  • only the third fastening means are constituted by a plurality of bolts, the first and second fixing means being each preferably constituted by a plurality of crimping elements.
  • the turbomachine of the invention further comprises a closing ferrule of ceramic composite material fixedly secured to said downstream end of said combustion chamber and intended to form a support plane for a seal ensuring the seal between said chamber of combustion and said dispenser.
  • This closure ferrule is preferably brazed to said downstream end of the combustion chamber. It may comprise a return portion disposed in the extension of the wall of the combustion chamber.
  • the support plane of the seal is formed in a plane perpendicular to said longitudinal axis of said combustion chamber.
  • the support plane of the seal is formed in a plane parallel to said longitudinal axis of said combustion chamber.
  • the seal is preferably of the "omega" seal type.
  • the support plane of the seal is formed in a plane inclined with respect to said longitudinal axis of said combustion chamber.
  • the seal is preferably of the type "slatted" seal held against said closure ferrule by means of an elastic element secured to said distributor.
  • this seal may comprise a plurality of calibrated leak holes.
  • Through-holes 54, 56 formed in the external metal 46 and inner metal platforms 48 of the distributor 42 are furthermore provided to ensure cooling of the fixed vanes 44 of the inlet distributor of the rotor of the high-pressure turbine from the compressed oxidant 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 fixing tabs are mounted for a first part of them (see the tab referenced 58) between the outer annular casing 12 and the outer axial wall 26 of the combustion chamber and for a second part (like the tongue 60). between the inner annular casing 14 and the inner axial wall 28 of the combustion chamber.
  • Each flexible fastening tab of metal material for example the tongue 58 shown on the figure 3 , consists of three branches connected in a star to have a general shape of Y with three points of attachment, the ends 62a, 62b; 64a, 64b of two of these three branches being fixed integrally at a downstream end, opposite to the injection system 20, flange 68, 70 (that is disposed in a plane perpendicular to the longitudinal axis 10 of the chamber ) of the outer axial wall 26 and inner 28 of the composite material combustion chamber by respectively first 72a, 74a and second 72b, 74b fixing means and the end 76; 78 of the last of these three branches being fastened integrally to one or other of the outer metal annular envelopes 12 or inner 14 by third fastening means 80, 82.
  • third fastening means 80, 82 Note that, depending on the required flexibility, a constant width or not, a form of tongue U or V, or other, since it has three points of attachment would be equally conceivable.
  • a ceramic composite closure ferrule 84, 86 is fixedly held, for example by brazing, against the flange 68, 70 of the combustion chamber to form a support plane for an omega-type circular seal.
  • 88, 90 mounted in a groove 92, 94 of each of the outer 46 and inner platforms 48 of the distributor and intended to seal between the combustion chamber 24 and the distributor 42.
  • this ferrule makes it possible, by its sufficient thickness to "drown" the screw heads of the first 72a, 74a and second 72b, 74b fixing means.
  • the figure 4 illustrates a second embodiment of the invention in which the downstream end of the combustion chamber no longer has a flange configuration, perpendicular to the longitudinal axis of the combustion chamber, but on the contrary a configuration is parallel to this axis is inclined with respect thereto (this inclination being able to go up to 90 °).
  • These non-perpendicular configurations of the downstream end of the chamber provide an improvement in the manufacture of the walls of the chamber, in particular by allowing better densification of the material in the rays.
  • the downstream end 70 of the internal axial wall 28 of the combustion chamber has a configuration parallel to the longitudinal axis 10 of this chamber (see the detail of FIG. figure 6 ) and comes in radial support, via the shell of composite material 86, against the inner circular platform 48 of the distributor.
  • this platform is provided with a groove 94 in which is housed a gasket 90 of "omega" type for sealing between the combustion chamber 24 and the distributor 42, at the inner axial wall of this chamber.
  • the downstream end 68 of the outer axial wall 26 of the combustion chamber has for its part, as shown in the detail of the figure 5 , a configuration inclined relative to the longitudinal axis 10 of the chamber.
  • a shell of composite material 84 is preferably brazed to the downstream end to form a support plane for a seal sealing between the combustion chamber 24 and the distributor 42, at this time of the outer axial wall of this chamber.
  • the seal now consists of a "lamellar" type circular seal 106 held against an elastic member 108, preferably a leaf spring, integral with the dispenser.
  • FIG. 5A illustrates another variant of an embodiment of the invention in which the fixing of the tongues 58 to the downstream end of the combustion chamber 68 is performed by a crimped connection, the bolts 72a, 72b being replaced by elements of crimp 72c, 72d.
  • the closure ring 84 is advantageously provided with a chamber return portion 84a disposed in the extension of the outer wall 26 of the combustion chamber.
  • calibrated leakage orifices 110 are provided at the joint 106.

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)
  • Gasket Seals (AREA)

Claims (13)

  1. Turbomaschine, die - in einem ringförmigen Gehäuse aus metallischem Material (12, 14) und entlang einer Strömungsrichtung F der Gase - eine Einheit zum Einspritzen eines Treibstoffs (20; 22), eine ringförmige Brennkammer aus Verbundwerkstoff (24) mit einer Längsachse (10) sowie ein ringförmiges Leitrad aus metallischem Material (42), welches die Eintrittsstufe mit Leitschaufeln (44) einer Hochdruckturbine bildet, umfaßt, wobei die Brennkammer aus Verbundwerkstoff in dem ringförmigen Metallgehäuse durch eine Vielzahl von flexiblen Metallzungen (58, 60), die um die Brennkammer gleichmäßig verteilt sind, in Position gehalten ist, dadurch gekennzeichnet, daß eine jede dieser Zungen drei sternförmig verbundene Schenkel aufweist, wobei die Enden (62a, 62b; 64a, 64b) von zwei dieser drei Schenkel mit einem zu dem Einspritzsystem (20) entgegengesetzten stromabwärtigen Ende (68, 70) der Brennkammer aus Verbundwerkstoff (26, 28) durch jeweils erste (72a, 72c; 74a) und zweite (72b, 72d; 74b) Befestigungsmittel fest verbunden sind und wobei das Ende (76, 78) des letzten dieser drei Schenkel mit dem ringförmigen Metallgehäuse (12, 14) durch dritte Befestigungsmittel (80, 82) fest verbunden ist, wobei die Flexibilität der Befestigungszungen bei hohen Temperaturen eine freie radiale Ausdehnung der Brennkammer aus Verbundwerkstoff gegenüber dem ringförmigen Metallgehäuse ermöglicht.
  2. Turbomaschine nach Anspruch 1, dadurch gekennzeichnet, daß die ersten, zweiten und dritten Befestigungsmittel jeweils von einer Vielzahl von Bolzen (72a, 74a; 72b, 74b; 80, 82) gebildet sind.
  3. Turbomaschine nach Anspruch 1, dadurch gekennzeichnet, daß die ersten und die zweiten Befestigungsmittel jeweils von einer Vielzahl von Crimpelementen (72c, 72d) gebildet sind, während die dritten Befestigungsmittel von einer Vielzahl von Bolzen (80, 82) gebildet sind.
  4. Turbomaschine nach Anspruch 1, dadurch gekennzeichnet, daß sie ferner einen Schließring aus keramischem Verbundwerkstoff (84, 86) umfaßt, der mit dem stromabwärtigen Ende der Brennkammer fest verbunden und dazu bestimmt ist, eine Auflageebene für eine Dichtung (88, 90, 106) zu bilden, welche die Dichtigkeit zwischen der Brennkammer und dem Leitrad sicherstellt.
  5. Turbomaschine nach Anspruch 5, dadurch gekennzeichnet, daß der Schließring an das stromabwärtige Ende der Brennkammer angelötet ist.
  6. Turbomaschine nach Anspruch 5, dadurch gekennzeichnet, daß der Schließring einen Rückwärtsteil (84a) aufweist, der in der Verlängerung der Wand der Brennkammer (26) angeordnet ist.
  7. Turbomaschine nach Anspruch 5, dadurch gekennzeichnet, daß die Auflageebene der Dichtung in einer zur Längsachse der Brennkammer senkrecht verlaufenden Ebene gebildet ist.
  8. Turbomaschine nach Anspruch 5, dadurch gekennzeichnet, daß die Auflageebene der Dichtung in einer zur Längsachse der Brennkammer parallel verlaufenden Ebene gebildet ist.
  9. Turbomaschine nach Anspruch 7 oder Anspruch 8, dadurch gekennzeichnet, daß die Dichtung vom Typ "Omega" Dichtung (88, 90) ist.
  10. Turbomaschine nach Anspruch 5, dadurch gekennzeichnet, daß die Auflageebene der Dichtung in einer gegenüber der Längsachse der Brennkammer geneigten Ebene gebildet ist.
  11. Turbomaschine nach Anspruch 10, dadurch gekennzeichnet, daß die Dichtung vom Typ "Lamellen"-Dichtung (106) ist.
  12. Turbomaschine nach Anspruch 11, dadurch gekennzeichnet, daß die "Lamellen"-Dichtung mittels eines mit dem Leitrad fest verbundenen elastischen Elements (108) an dem Schließring gehalten ist.
  13. Turbomaschine nach Anspruch 11, dadurch gekennzeichnet, daß die "Lamellen"-Dichtung eine Vielzahl kalibrierter Leckageöffnungen (110) aufweist.
EP02291365A 2001-06-06 2002-06-04 Elastische Befestigung einer Turbinenbrennkammer aus keramischem Matrix-Verbundwerkstoff in einem metallischen Gehäuse Expired - Lifetime EP1265036B1 (de)

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
FR0107361 2001-06-06
FR0107361A FR2825781B1 (fr) 2001-06-06 2001-06-06 Montage elastique de chambre ce combustion cmc de turbomachine dans un carter metallique

Publications (2)

Publication Number Publication Date
EP1265036A1 EP1265036A1 (de) 2002-12-11
EP1265036B1 true EP1265036B1 (de) 2008-10-22

Family

ID=8863985

Family Applications (1)

Application Number Title Priority Date Filing Date
EP02291365A Expired - Lifetime EP1265036B1 (de) 2001-06-06 2002-06-04 Elastische Befestigung einer Turbinenbrennkammer aus keramischem Matrix-Verbundwerkstoff in einem metallischen Gehäuse

Country Status (5)

Country Link
US (1) US6732532B2 (de)
EP (1) EP1265036B1 (de)
JP (1) JP4031292B2 (de)
DE (1) DE60229466D1 (de)
FR (1) FR2825781B1 (de)

Cited By (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN101981382B (zh) * 2008-03-28 2012-06-13 三菱重工业株式会社 燃气轮机及燃气轮机的燃烧器插入孔形成方法

Families Citing this family (24)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
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
US6775985B2 (en) * 2003-01-14 2004-08-17 General Electric Company Support assembly for 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
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
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
GB2422874A (en) * 2005-02-05 2006-08-09 Alstom Technology Ltd Gas turbine burner expansion bar structure
US7546743B2 (en) * 2005-10-12 2009-06-16 General Electric Company Bolting configuration for joining ceramic combustor liner to metal mounting attachments
FR2892181B1 (fr) * 2005-10-18 2008-02-01 Snecma Sa Fixation d'une chambre de combustion a l'interieur de son carter
US7637110B2 (en) * 2005-11-30 2009-12-29 General Electric Company Methods and apparatuses for assembling a gas turbine engine
US7493771B2 (en) * 2005-11-30 2009-02-24 General Electric Company Methods and apparatuses for assembling a gas turbine engine
US7523616B2 (en) * 2005-11-30 2009-04-28 General Electric Company Methods and apparatuses for assembling a gas turbine engine
FR2897418B1 (fr) 2006-02-10 2013-03-01 Snecma Chambre de combustion annulaire d'une turbomachine
US8863528B2 (en) * 2006-07-27 2014-10-21 United Technologies Corporation Ceramic combustor can for a gas turbine engine
FR2919380B1 (fr) * 2007-07-26 2013-10-25 Snecma Chambre de combustion d'une turbomachine.
US20090067917A1 (en) * 2007-09-07 2009-03-12 The Boeing Company Bipod Flexure Ring
US20120047909A1 (en) * 2010-08-24 2012-03-01 Nuovo Pignone S.P.A. Combustor liner concentric support and method
US8919134B2 (en) * 2011-01-26 2014-12-30 United Technologies Corporation Intershaft seal with support linkage
US9335051B2 (en) * 2011-07-13 2016-05-10 United Technologies Corporation Ceramic matrix composite combustor vane ring assembly
EP2841749B1 (de) 2012-04-27 2020-02-26 General Electric Company Verbindung von ringförmigen gasturbinenmotorelementen
US9435266B2 (en) 2013-03-15 2016-09-06 Rolls-Royce North American Technologies, Inc. Seals for a gas turbine engine
US10436446B2 (en) 2013-09-11 2019-10-08 General Electric Company Spring loaded and sealed ceramic matrix composite combustor liner
US9404421B2 (en) 2014-01-23 2016-08-02 Siemens Energy, Inc. Structural support bracket for gas flow path
US20170059165A1 (en) 2015-08-28 2017-03-02 Rolls-Royce High Temperature Composites Inc. Cmc cross-over tube

Family Cites Families (11)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2268464A (en) * 1939-09-29 1941-12-30 Bbc Brown Boveri & Cie Combustion chamber
US2509503A (en) * 1946-02-12 1950-05-30 Lucas Ltd Joseph Combustion chamber for prime movers
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.
US4688378A (en) * 1983-12-12 1987-08-25 United Technologies Corporation One piece band seal
US4821522A (en) * 1987-07-02 1989-04-18 United Technologies Corporation Sealing and cooling arrangement for combustor vane interface
US5291733A (en) * 1993-02-08 1994-03-08 General Electric Company Liner mounting assembly
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

Cited By (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN101981382B (zh) * 2008-03-28 2012-06-13 三菱重工业株式会社 燃气轮机及燃气轮机的燃烧器插入孔形成方法

Also Published As

Publication number Publication date
US20020184890A1 (en) 2002-12-12
DE60229466D1 (de) 2008-12-04
JP4031292B2 (ja) 2008-01-09
FR2825781A1 (fr) 2002-12-13
FR2825781B1 (fr) 2004-02-06
JP2003021334A (ja) 2003-01-24
EP1265036A1 (de) 2002-12-11
US6732532B2 (en) 2004-05-11

Similar Documents

Publication Publication Date Title
EP1265036B1 (de) Elastische Befestigung einer Turbinenbrennkammer aus keramischem Matrix-Verbundwerkstoff in einem metallischen Gehäuse
EP1265034B1 (de) Befestigung einer Turbinenbrennkammer aus keramischem Matrix-Verbundswerkstoff mit gelöteten Befestigungsfüssen
EP1265035B1 (de) Doppelbefestigung einer Turbinenbrennkammer aus keramischem Matrix-Verbundwerkstoff
EP1265037B1 (de) Verdünnungsluftlöcher verwendende Befestigung einer Turbinenbrennkammer aus keramischem Matrix-Verbundwerkstoff
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
EP1734305B1 (de) Ringförmige Brennkammer für eine Turbine
CA2577514C (fr) Chambre de combustion annulaire d'une turbomachine
EP1265033B1 (de) Brennkammer mit einer Befestigungsanordnung der Kammerendwand
EP1479975B1 (de) Brennkammer mit einer flexiblen Verbindung zwischen einem Kammerboden und einer Kammerwand
EP1705342B1 (de) Vorrichtung zur Verbindung eines Kühlluftplenums mit einer Statorschaufel in einer Turbomaschine
EP2917519B1 (de) Aufnahme eines entlüftungsrohr einer turbomachine
EP1519009B1 (de) Turbomaschine mit Luftentnahme für die Kabine über ein Kugelgelenkrohr
FR2896575A1 (fr) Chambre de combustion annulaire d'une turbomachine
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
FR2903151A1 (fr) Dispositif de ventilation d'un carter d'echappement dans une turbomachine
EP3721058B1 (de) Verbindung zwischen einem leitkranzsektor aus verbundwerkstoff mit keramischer matrix und einem metallischen träger einer turbine einer turbomaschine
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

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: 60229466

Country of ref document: DE

Date of ref document: 20081204

Kind code of ref document: P

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: 20090723

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: DE

Payment date: 20210519

Year of fee payment: 20

Ref country code: FR

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: 60229466

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