EP1265032B1 - Gasturbinenbrennkammer aus Verbundwerkstoff mit keramischer Matrix - Google Patents

Gasturbinenbrennkammer aus Verbundwerkstoff mit keramischer Matrix Download PDF

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Publication number
EP1265032B1
EP1265032B1 EP02291361A EP02291361A EP1265032B1 EP 1265032 B1 EP1265032 B1 EP 1265032B1 EP 02291361 A EP02291361 A EP 02291361A EP 02291361 A EP02291361 A EP 02291361A EP 1265032 B1 EP1265032 B1 EP 1265032B1
Authority
EP
European Patent Office
Prior art keywords
combustion chamber
flange
nozzle
shell
turbomachine according
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
EP02291361A
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English (en)
French (fr)
Other versions
EP1265032A1 (de
Inventor
Alexandre Forestier
Didier Hernandez
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 Moteurs SA
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 Moteurs SA filed Critical SNECMA Moteurs SA
Publication of EP1265032A1 publication Critical patent/EP1265032A1/de
Application granted granted Critical
Publication of EP1265032B1 publication Critical patent/EP1265032B1/de
Anticipated expiration legal-status Critical
Expired - Lifetime legal-status Critical Current

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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
    • F23R3/007—Continuous combustion chambers using liquid or gaseous fuel constructed mainly of ceramic components
    • 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
    • F23R3/50—Combustion chambers comprising an annular flame tube within an annular casing
    • 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
    • F23M—CASINGS, LININGS, WALLS OR DOORS SPECIALLY ADAPTED FOR COMBUSTION CHAMBERS, e.g. FIREBRIDGES; DEVICES FOR DEFLECTING AIR, FLAMES OR COMBUSTION PRODUCTS IN COMBUSTION CHAMBERS; SAFETY ARRANGEMENTS SPECIALLY ADAPTED FOR COMBUSTION APPARATUS; DETAILS OF COMBUSTION CHAMBERS, NOT OTHERWISE PROVIDED FOR
    • F23M2900/00—Special features of, or arrangements for combustion chambers
    • F23M2900/05002—Means for accommodate thermal expansion of the wall liner

Definitions

  • the present invention relates to the field of turbomachines and more particularly it relates to the interface between the high pressure turbine and the combustion chamber of turbojets having a combustion chamber CMC (ceramic matrix composite).
  • CMC ceramic matrix composite
  • the high pressure turbine including its inlet distributor (HPT nozzle), the combustion chamber as well as the casing (also called envelope) of this room are made in a same material, usually of metal type.
  • HPT nozzle inlet distributor
  • the combustion chamber as well as the casing (also called envelope) of this room are made in a same material, usually of metal type.
  • CMC-type high temperature composite materials as shown in US-A-5,291,732.
  • difficulties of implementation of these materials and their cost make that their use is most often limited to the combustion chamber itself, the distributor inlet of the high pressure turbine and the remaining housing then made more typically in metallic materials.
  • metallic materials and Composite materials have very different coefficients of thermal expansion. This results in particularly acute interface problems at the level of the distributor, inlet of the high pressure turbine, and connection with the housing of the bedroom.
  • the present invention overcomes these disadvantages by proposing a connection crankcase with the capacity to absorb the displacements induced by differences in the expansion coefficients of these parts.
  • An object of the invention is also to propose a structure of simple form and whose manufacture is particularly easy.
  • a turbomachine comprising, in a envelope of metallic material and in a direction F of gas flow, a fuel injection assembly, combustion chamber made of material composite and a metal material distributor forming the entrance stage to vanes of a high-pressure turbine, and said distributor being supported by said envelope and fixed thereto by first removable fixing means, characterized in that said combustion chamber is mounted floating in said envelope and held in position by the only said distributor to which it is resiliently secured by second removable attachment means.
  • the first removable fastening means are adapted to allow radial free expansion of said dispenser with respect to said envelope.
  • said second means of removable fasteners comprise on the one hand first holding means for pinch an end internal axial wall of said chamber of combustion between an internal circular platform of the dispenser and a flange supporting an inner annular wall of said envelope and second holding means for holding a wall with elastic prestressing external axial end of said combustion chamber on a platform external circular of the distributor.
  • the support flange is sectorised to compensate for the differences circumferential geometries resulting from the existing differential expansion at high temperatures between said inner circular distributor platform and said inner axial wall of the combustion chamber.
  • This support flange is mounted between a flange of said inner annular wall of the envelope and a ferrule made of metallic material held against this flange by said first removable fastening means.
  • said first removable fastening means have a plurality of bolts, each of which has screw axes running through a corresponding oblong hole of said support flange is provided with a shoulder on which is pressed against said ferrule so as to allow sliding of said support flange between said ferrule and said flange of the inner annular wall of the envelope.
  • said flange of the wall inner ring of the envelope has a circular groove for receiving a "omega" type circular sealing gasket intended to seal between said flange of the inner annular wall of the casing and said support flange.
  • a composite material shell advantageously brazed to said outer end wall of the combustion chamber is maintained with a elastic prestress against said external circular distributor platform by the second holding means, said ferrule having a groove to receive an omega-type circular seal intended for sealing between said outer end wall of the chamber of combustion and said outer circular platform of the dispenser.
  • this space 16 comprising, in the direction of gas flow, firstly an injection assembly formed of a plurality of injection systems 20 and regularly distributed around the duct 18 and each having a nozzle fuel injection device 22 fixed to the outer annular casing 12 (in a simplification of the drawings the mixer and the baffle associated with each injection nozzle were not shown), then a combustion chamber in high temperature composite material 24, CMC type or other (carbon by example), formed of an outer axial wall 26 and an inner axial wall 28, both coaxial axis 10, and a transverse wall 30 which constitutes the bottom of this combustion chamber and which has flaps 32, 34 fixed by all suitable means, for example metal or refractory bolts with screws conical), on the upstream ends 36, 38 of the axial walls 26, 28, this bottom of the chamber 30 being provided with orifices 40 to allow fuel injection and part of the oxidant in the combustion chamber 24, and finally a annular distributor 42 in metallic material forming an input stage of a high pressure turbine (not shown) and conventionally comprising a plurality of fixed vanes 44
  • the combustion chamber is mounted floating in the annular envelope and held in position by the only distributor to which it is resiliently secured by second removable attachment means which comprise on the one hand first holding means 52 for holding by pinching an end internal axial wall 54 of the combustion chamber (opposite the upstream end 38) between the inner circular platform of the distributor 48 and a flange 56 serving to support the inner annular envelope 14 and second holding means 58 for holding with prestressing elastic 60 an end outer axial wall 62 of said chamber of combustion (opposite the upstream end 36) on the outer circular platform of the dispenser 46.
  • the support flange 56 is mounted between a flange 64 of the inner annular casing 14 and a ferrule made of metallic material 66 held against this flange by the first removable fastening means 50.
  • Through orifices 68, 70 for the passage of the compressed oxidant previously separated, at the outlet of the diffusion duct 18, in at least two streams F1, F2 separate flowing from both sides of the combustion chamber 24 (and notably ensuring its cooling), are provided in the platforms external metal 46 and internal 48 of the distributor 42 to ensure a cooling the vanes fixed 44 of the distributor at the inlet of the rotor of the turbine high pressure.
  • the combustion chamber 24 having a coefficient of thermal expansion very different from other metal parts forming the turbomachine, in particular the dispenser 42 to which it is attached and the annular casing 12, 14, it is provided that the first removable fixing means 50 are adapted to allow at high temperatures radial free expansion of the dispenser with respect to the annular envelope.
  • the support flange 56 is pierced with holes oblongs 72 for cooperating with the screw axes of the plurality of bolts 50 a shoulder 74 serves to support the ferrule 66 so as to allow a sliding of this support flange between the ferrule and the flange 64 of the envelope 14.
  • this flange is sectorised to compensate for differences circumferential geometries resulting from the differential expansion existing at these high temperatures between the inner circular platform 48 of the dispenser and the internal axial wall 28, 54 of the combustion chamber.
  • the flange of the inner annular casing 64 comprises a circular groove 76 for receiving a circular seal of type "Omega" 78 for sealing between this flange of the envelope ring and the support flange 56.
  • the outer circular platform distributor 46 comprises a flange 80 provided with a circular groove 82 for receive a plate seal 84, one end of which will come into contact with the outer annular casing 12 to seal against the flow F1.
  • the seal between the combustion chamber 24 and the distributor 42 it is provided between the outer end wall of the combustion chamber 62 and the outer circular platform of the distributor 46 also by means of a omega-type seal ring 86 mounted in a groove circular 88 of a shell of composite material 90, preferably brazed on the outer end wall of the combustion chamber 62, and maintained with an elastic prestress (obtained for example by the spring 60) against the platform external circular distributor 46 by the second holding means 58.

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)

Claims (9)

  1. Turbomaschine, umfassend in einer Hülle aus metallischem Material (12, 14) und entlang einer Gasabflussrichtung F eine Einspritzeinheit für einen Treibstoff (20; 22), eine Brennkammer aus Verbundmaterial (24) und einen Verteiler aus metallischem Material (42), der die Eingangsstufe mit festen Schaufeln (44) einer Hochdruckturbine bildet, wobei der Verteiler von der Hülle getragen wird und an dieser durch erste abnehmbare Befestigungsmittel (50) befestigt ist, dadurch gekennzeichnet, dass die Brennkammer schwimmend in der Hülle angeordnet und durch den sogenannten einzigen Verteiler in Position gehalten wird, mit dem sie elastisch durch zweite abnehmbare Befestigungsmittel (52, 58, 60) verbunden ist.
  2. Turbomaschine nach Anspruch 1, dadurch gekennzeichnet, dass die ersten abnehmbaren Befestigungsmittel derart ausgeführt sind, dass sie eine freie Radialdehnung des Verteilers in Bezug auf die Hülle ermöglichen.
  3. Turbomaschine nach Anspruch 1, dadurch gekennzeichnet, dass die zweiten abnehmbaren Befestigungsmittel zu einem Teil erste Haltemittel (52), um durch Klemmen eine innere axiale Endwand (54) der Brennkammer zwischen einer inneren kreisförmigen Plattform des Verteilers (48) und einem Flansch (56) zu halten, der als Stütze für eine innere ringförmige Wand (14) der Hülle dient, und zweite Haltemittel (58) umfassen, um mit einer elastischen Vorspannung (60) eine äußere axiale Endwand (62) der Brennkammer auf einer äußeren kreisförmigen Plattform des Verteilers (46) zu halten.
  4. Turbomaschine nach Anspruch 3, dadurch gekennzeichnet, dass der Stützflansch in Abschnitte unterteilt ist, um die geometrischen Umfangsabweichungen auszugleichen, die sich aus der Differentialdehnung ergeben, die bei hohen Temperaturen zwischen der inneren kreisförmigen Plattform des Verteilers und der inneren axialen Wand der Brennkammer vorhanden ist.
  5. Turbomaschine nach Anspruch 3, dadurch gekennzeichnet, dass der Stützflansch zwischen einem Flansch (64) der inneren ringförmigen Wand der Hülle und einem Reif aus metallischem Material (66) montiert ist, der gegen diesen Flansch durch die ersten abnehmbaren Befestigungsmittel gehalten wird.
  6. Turbomaschine nach Anspruch 5, dadurch gekennzeichnet, dass die ersten abnehmbaren Befestigungsmittel eine Vielzahl von Bolzen (50) umfassen, wobei jede deren Schraubachsen, die durch ein entsprechendes Langloch (72) des Flansches hindurchgeführt wird, mit einem Absatz (74) versehen ist, auf dem der Reif derart abgestützt ist, dass ein Gleiten des Stützflansches zwischen dem Reif und dem Flansch der inneren ringförmigen Wand der Hülle möglich ist.
  7. Turbomaschine nach Anspruch 6, dadurch gekennzeichnet, dass der Flansch der inneren ringförmigen Wand der Hülle eine kreisförmige Nut (76) umfasst, um eine kreisförmige Dichtung vom Typ "Omega" (78) aufzunehmen, die dazu bestimmt ist, die Dichtigkeit zwischen dem Flansch der inneren ringförmigen Wand der Hülle und dem Stützflansch sicher zu stellen.
  8. Turbomaschine nach Anspruch 3, dadurch gekennzeichnet, dass sie ferner einen Reif aus Verbundmaterial (90) umfasst, der vorzugsweise auf die äußere Endwand der Brennkammer gelötet ist und mit einer elastischen Vorspannung (60) an der äußeren kreisförmigen Plattform des Verteilers durch die zweiten Haltemittel gehalten wird.
  9. Turbomaschine nach Anspruch 8, dadurch gekennzeichnet, dass der Reif eine kreisförmige Nut (88) umfasst, um eine kreisförmige Dichtung vom Typ "Omega" (86) aufzunehmen, die dazu bestimmt ist, die Dichtigkeit zwischen der äußeren Endwand der Brennkammer und der äußeren kreisförmigen Plattform des Verteilers sicher zu stellen.
EP02291361A 2001-06-06 2002-06-04 Gasturbinenbrennkammer aus Verbundwerkstoff mit keramischer Matrix Expired - Lifetime EP1265032B1 (de)

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
FR0107360A FR2825780B1 (fr) 2001-06-06 2001-06-06 Architecure de chambre de combustion de turbomachine en materiau a matrice ceramique
FR0107360 2001-06-06

Publications (2)

Publication Number Publication Date
EP1265032A1 EP1265032A1 (de) 2002-12-11
EP1265032B1 true EP1265032B1 (de) 2004-10-06

Family

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Family Applications (1)

Application Number Title Priority Date Filing Date
EP02291361A Expired - Lifetime EP1265032B1 (de) 2001-06-06 2002-06-04 Gasturbinenbrennkammer aus Verbundwerkstoff mit keramischer Matrix

Country Status (5)

Country Link
US (1) US6679062B2 (de)
EP (1) EP1265032B1 (de)
JP (1) JP3983603B2 (de)
DE (1) DE60201467T2 (de)
FR (1) FR2825780B1 (de)

Families Citing this family (27)

* 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
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
US7152411B2 (en) * 2003-06-27 2006-12-26 General Electric Company Rabbet mounted combuster
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
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
FR2871844B1 (fr) * 2004-06-17 2006-09-29 Snecma Moteurs Sa Montage etanche d'un distributeur de turbine haute pression sur une extremite d'une chambre de combustion dans une turbine a gaz
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
US7647779B2 (en) 2005-04-27 2010-01-19 United Technologies Corporation Compliant metal support for ceramic combustor liner in a gas turbine engine
US20070134087A1 (en) * 2005-12-08 2007-06-14 General Electric Company Methods and apparatus for assembling turbine engines
US8038389B2 (en) 2006-01-04 2011-10-18 General Electric Company Method and apparatus for assembling turbine nozzle assembly
EP1843009A1 (de) 2006-04-06 2007-10-10 Siemens Aktiengesellschaft Leitschaufelsegment einer thermischen Strömungsmaschine, zugehöriges Herstellungsverfahren sowie thermische Strömungsmaschine
FR2906350B1 (fr) 2006-09-22 2009-03-20 Snecma Sa Chambre de combustion annulaire d'une turbomachine
DE102006060857B4 (de) * 2006-12-22 2014-02-13 Deutsches Zentrum für Luft- und Raumfahrt e.V. CMC-Brennkammerauskleidung in Doppelschichtbauweise
EP1985806A1 (de) * 2007-04-27 2008-10-29 Siemens Aktiengesellschaft Deckbandkühlung einer Turbinenleitschaufel
US8745989B2 (en) 2009-04-09 2014-06-10 Pratt & Whitney Canada Corp. Reverse flow ceramic matrix composite combustor
US8206096B2 (en) * 2009-07-08 2012-06-26 General Electric Company Composite turbine nozzle
US8226361B2 (en) * 2009-07-08 2012-07-24 General Electric Company Composite article and support frame assembly
US9234431B2 (en) * 2010-07-20 2016-01-12 Siemens Energy, Inc. Seal assembly for controlling fluid flow
FR2963061B1 (fr) * 2010-07-26 2012-07-27 Snecma Systeme d?injection de carburant pour turbo-reacteur et procede d?assemblage d?un tel systeme d?injection
US9290261B2 (en) 2011-06-09 2016-03-22 United Technologies Corporation Method and assembly for attaching components
US10059431B2 (en) 2011-06-09 2018-08-28 United Technologies Corporation Method and apparatus for attaching components having dissimilar rates of thermal expansion
US9335051B2 (en) * 2011-07-13 2016-05-10 United Technologies Corporation Ceramic matrix composite combustor vane ring assembly
WO2014149110A2 (en) 2013-03-15 2014-09-25 Sutterfield David L Seals for a gas turbine engine
GB201315871D0 (en) * 2013-09-06 2013-10-23 Rolls Royce Plc A combustion chamber arrangement
EP3236155B1 (de) * 2016-04-22 2020-05-06 Rolls-Royce plc Brennkammer mit segmentierter wand
US10495001B2 (en) 2017-06-15 2019-12-03 General Electric Company Combustion section heat transfer system for a propulsion system
US11708765B1 (en) 2022-05-13 2023-07-25 Raytheon Technologies Corporation Gas turbine engine article with branched flange

Family Cites Families (6)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3775975A (en) * 1972-09-05 1973-12-04 Gen Electric Fuel distribution system
US4912922A (en) * 1972-12-19 1990-04-03 General Electric Company Combustion chamber construction
US3965066A (en) * 1974-03-15 1976-06-22 General Electric Company Combustor-turbine nozzle interconnection
US5335502A (en) * 1992-09-09 1994-08-09 General Electric Company Arched combustor
US5291732A (en) * 1993-02-08 1994-03-08 General Electric Company Combustor liner support assembly
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

Also Published As

Publication number Publication date
US20020184891A1 (en) 2002-12-12
DE60201467D1 (de) 2004-11-11
FR2825780B1 (fr) 2003-08-29
EP1265032A1 (de) 2002-12-11
JP3983603B2 (ja) 2007-09-26
JP2003035104A (ja) 2003-02-07
US6679062B2 (en) 2004-01-20
DE60201467T2 (de) 2006-03-09
FR2825780A1 (fr) 2002-12-13

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