EP1956297B1 - Gasturbinenverbrennungskammer - Google Patents

Gasturbinenverbrennungskammer Download PDF

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Publication number
EP1956297B1
EP1956297B1 EP08075032.6A EP08075032A EP1956297B1 EP 1956297 B1 EP1956297 B1 EP 1956297B1 EP 08075032 A EP08075032 A EP 08075032A EP 1956297 B1 EP1956297 B1 EP 1956297B1
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EP
European Patent Office
Prior art keywords
annular
fairing
combustion chamber
chamber
internal
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.)
Active
Application number
EP08075032.6A
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English (en)
French (fr)
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EP1956297A1 (de
Inventor
Mario César De Sousa
Morgan Robin
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
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SNECMA SAS
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Publication of EP1956297A1 publication Critical patent/EP1956297A1/de
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    • 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
    • F23R3/28—Continuous combustion chambers using liquid or gaseous fuel characterised by the fuel supply
    • F23R3/283—Attaching or cooling of fuel injecting means including supports for fuel injectors, stems, or lances
    • 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
    • F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
    • F23R2900/00—Special features of, or arrangements for continuous combustion chambers; Combustion processes therefor
    • F23R2900/00005—Preventing fatigue failures or reducing mechanical stress in gas turbine 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
    • F23R2900/00—Special features of, or arrangements for continuous combustion chambers; Combustion processes therefor
    • F23R2900/00017—Assembling combustion chamber liners or subparts

Definitions

  • the present invention relates to an annular combustion chamber of a turbomachine, such as an airplane turbojet or turboprop.
  • An annular combustion chamber of a turbomachine comprises two coaxial cylindrical walls, connected at their upstream ends to a very rigid annular bottom wall of the chamber and having, at their downstream ends, fastening flanges on casings of the turbomachine. It also comprises an upstream annular fairing fixed to the chamber bottom and intended to guide the flow of air entering or bypassing the combustion chamber.
  • the assembly of the upstream part of the combustion chamber is achieved by superposition of the radially inner and outer ends of the fairing on the radially inner and outer ends respectively of the cylindrical walls of the chamber, the assembly being fixed by bolting or welding on radially inner and outer annular flanges respectively of the chamber bottom.
  • Bolting fastening is generally preferred since the maintenance operations performed on the combustion chamber are simpler and less expensive than for welding attachment.
  • Such a fixing technique is described by the document GB2263733 .
  • the parts of the fairing, the superimposed ends of the cylindrical walls and the chamber bottom have complementary corrugated surfaces, the fasteners being made at the vertices of the corrugations.
  • the present invention relates to a combustion chamber for a turbomachine, which avoids the aforementioned drawbacks of the prior art simply, efficiently and economically.
  • annular combustion chamber of a turbomachine comprising two cylindrical walls radially inner and radially outer respectively relative to the axis of the turbomachine, fixed by bolting at their upstream ends on an inner annular flange and on a outer annular flange of an annular chamber bottom, and an annular fairing extending upstream from the chamber bottom, characterized in that the inner and outer downstream annular ends of the fairing are bolted to the inner annular flanges and outer respectively of the chamber bottom, in axial alignment with the annular ends of the inner and outer walls of the chamber.
  • the assembly of the upstream end of the combustion chamber is thus done by radial superposition of two parts and not more than three parts which reduces the accumulation of stiffness and the accumulation of manufacturing tolerances.
  • the tightening torque to be applied to the bolts can be optimized and the radial deformations of the chamber during the fixing of the fairing and the walls respectively on the bottom wall are reduced.
  • the aligned ends of the fairing and the cylindrical walls of the combustion chamber comprise indentations or complementary corrugations engaged in each other and traversed by fixing bolts on the chamber bottom.
  • the indentations or corrugations of the ends of the fairing and the cylindrical walls comprise an alternation of solid parts and hollow parts, the fastening bolts passing through the solid parts and being distributed in an annular row on the outer annular end of the shroud and on the corresponding end of the outer wall of the chamber, and in an annular row on the inner annular end of the shroud and on the corresponding end of the wall internal of the room.
  • the fixing bolts of the outer annular end of the fairing and the annular end of the outer wall are offset angularly with respect to the fixing bolts of the inner annular end of the fairing and the annular end of the wall. internal.
  • the configuration is such that a fixing bolt of the outer downstream end of the shroud is aligned radially not with a bolt for fixing the inner downstream end of the shroud.
  • each solid part of the indentations or corrugations comprises a single hole for the passage of fixing bolts.
  • the solid portions of the indentations or corrugations of the ends of the fairing comprise the same number of fixing bolts as the solid parts of the indentations or undulations of the ends of the walls of the chamber.
  • the solid portions of the indentations or corrugations of the ends of the fairing comprise a number of fastening bolts different from that of the solid parts of the indentations or corrugations of the ends of the walls of the chamber.
  • the annular fairing can be made in one piece, or in two radially inner and outer annular parts respectively.
  • the invention also relates to a turbomachine, such as an airplane turbojet or turboprop, characterized in that it comprises an annular combustion chamber of the type described above.
  • FIG 1 is a schematic half-view of an annular combustion chamber 10 according to the prior art to the invention, seen in section along the axis of rotation 12 of the turbomachine.
  • the combustion chamber 10 is supplied with air by a diffuser 14 mounted at the outlet of a high-pressure compressor 16. It comprises a radially inner cylindrical wall 18 and a radially outer cylindrical wall 20 connected upstream to an annular chamber bottom 22 and downstream to housings 24 and 26 via an inner annular flange 28 and an outer annular flange 30, respectively.
  • the chamber bottom 22 has openings 36 for the passage of air from the diffuser 14 and the fuel sprayed by injectors 34 carried by the outer casing 26.
  • Each injector 34 comprises a head 38 mounted on the chamber bottom and aligned with the axis 40 of an orifice 36.
  • An annular fairing 60 which extends upstream and which comprises orifices 44 for air passage and passage of the injectors, is fixed on the edges of the chamber bottom 22 with the ends of the cylindrical walls 18 and 20 of the combustion chamber.
  • the assembly of the upstream portion of the combustion chamber is made by interposing the inner ends 46 and outer 48 of the cylindrical walls between, on the one hand, the inner annular ends 50 and outer 52 of the shroud and on the other hand , the inner and outer annular rims 54 and 56 of the chamber bottom. These three pieces thus superimposed are fixed together by bolts 42, which results in a plurality of manufacturing tolerances and a plurality of stiffness.
  • the assembly of the upstream part of the chamber is thus achieved by radial superposition of two pieces and not more than three pieces. Consequently, the impact of the accumulations of the manufacturing tolerances and the respective stiffnesses of the fairing, the walls and the bottom is lower, which facilitates the assembly of the chamber and improves the mechanical strength of the assembly.
  • the tightening torque to be applied to the fairing fastening bolts on the bottom wall flanges can be optimized by taking into account only the fairing and bottom wall stiffness and manufacturing tolerances. Similarly, for fixing the cylindrical walls on the rims of the chamber bottom, only the stiffness and manufacturing tolerances of the walls and the chamber bottom are taken into account. This assembly makes it possible to limit the radial deformations of the fairing and the cylindrical walls and to avoid the formation of additional air leaks that disrupt combustion and air flow.
  • the upstream ends of the inner cylindrical walls 18 and outer 20 comprise corrugations or indentations formed by an alternation of hollow portions 62 and solid portions 48 which extend in alignment with these walls.
  • the inner 50 and outer 52 downstream ends of the fairing comprise corrugations formed by an alternation of hollow portions 64 and solid portions 50.
  • the hollow portions 62 and the solid portions 48 of the cylindrical walls are engaged in the solid portions 50 and the hollow portions 64, respectively, of the annular fairing.
  • the fixing bolts on the chamber bottom pass through the solid portions of the corrugations and are distributed along an outer annular row and an inner annular row.
  • the outer annular row is formed by an alternation of bolts 66 for fixing the outer cylindrical wall on the flange 56 of the chamber bottom and bolts 68 for fixing the outer upstream annular end of the fairing on this flange.
  • the internal annular row of bolts is formed by alternating bolts 70 for fixing the inner cylindrical wall and bolts 72 for fixing the inner upstream annular end of the fairing on the rim 54 of the chamber bottom.
  • the bolts 68 for fixing the outer annular end 52 of the fairing are angularly offset by one pitch relative to the bolts 72 for fixing the inner annular end of the fairing, and the bolts 66 and 72, as well as the bolts 68 and 70 are aligned radially.
  • This method of fixing with angular offset presents the advantage of stiffening the entire combustion chamber, avoiding the formation of deformation lines between an internal bolt 72 and an external bolt 68 which would be diametrically opposed.
  • the frequency levels of the eigen modes of vibration are thus higher which makes it possible to eliminate the risks of propagation of cracks under the effect of the vibrations.
  • each solid part of the indentations or corrugations comprises a single orifice for the passage of a fixing bolt.
  • the solid portions of the corrugations of the ends of the fairing comprise either the same number, for example equal to 2, or a different number of fixing bolts than the solid parts of the corrugations of the ends of the walls of the chamber. .
  • the annular fairing can be made in one piece or in two radially inner and radially outer annular parts.
  • the invention is not limited to the previously described combustion chambers and is generally applicable to all types of combustion chambers, such as, for example, those adapted to receive a plurality of concentric ring nozzle heads. .

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  • Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)

Claims (9)

  1. Ringförmige Verbrennungskammer (10) einer Turbomaschine, umfassend zwei zylindrische, radial interne (18) und radial externe (20) Wände jeweils im Verhältnis zur Achse der Turbomaschine (12), die per Verbolzung (42) an ihren vorgeschalteten Enden auf einem ringförmigen internen Rand (54) und auf einem ringförmigen externen Rand (56) eines ringförmigen Kammerbodens (22) befestigt sind, und ein ringförmiges Gehäuse (60), das sich von dem Boden der Kammer nach oben erstreckt, dadurch gekennzeichnet, dass die ringförmigen nachgeschalteten internen (50) und externen (52) Enden des Gehäuses (60) per Verbolzung auf den ringförmigen internen (54) und externen (56) Rändern jeweils des Bodens der Kammer (22) in axialer Fluchtung der ringförmigen Enden der internen (46) und der externen (48) Wand der Kammer befestigt sind.
  2. Verbrennungskammer gemäß Anspruch 1, dadurch gekennzeichnet, dass die gefluchteten Enden des Gehäuses und der zylindrischen Wände der Kammer komplementäre Zahnungen oder Wellungen umfassen, die ineinander eingreifen und durch Befestigungsbolzen (66, 68, 70, 72) auf dem Boden der Kammer durchquert sind.
  3. Verbrennungskammer gemäß Anspruch 2, dadurch gekennzeichnet, dass die Zahnungen oder Wellungen ein Abwechseln von massiven Teilen und hohlen Teilen umfassen, wobei die Befestigungsbolzen die massiven Teile durchqueren und in einer ringförmigen Reihe auf dem ringförmigen externen Ende (52) des Gehäuses und auf dem entsprechenden Ende der externen Wand (48) der Kammer und in einer ringförmigen Reihe am ringförmigen internen Ende (50) des Gehäuses und auf dem entsprechenden Ende der internen Wand (46) der Kammer verteilt sind.
  4. Verbrennungskammer gemäß Anspruch 1 bis 3, dadurch gekennzeichnet, dass die Befestigungsbolzen des ringförmigen externen Endes (52) des Gehäuses und des ringförmigen Endes (48) der externen Wand winkelförmig im Verhältnis zu den Befestigungsbolzen des ringförmigen interne Endes (50) des Gehäuses und des ringförmigen Endes (46) der internen Wand versetzt sind.
  5. Verbrennungskammer gemäß Anspruch 3 oder 4, dadurch gekennzeichnet, dass jeder massive Teile der Zahnungen oder Wellungen eine einzige Befestigungsbolzen-Durchgangsöffnung umfasst.
  6. Verbrennungskammer gemäß Anspruch 3 oder 4, dadurch gekennzeichnet, dass die massiven Teile der Zahnungen oder Wellungen der Enden des Gehäuses dieselbe Anzahl von Befestigungsbolzen umfassen wie die massiven Teile der Zahnungen oder Wellungen der Enden der Wände der Kammer.
  7. Verbrennungskammer gemäß Anspruch 3 oder 4, dadurch gekennzeichnet, dass die massiven Teile der Zahnungen oder Wellungen der Enden des Gehäuses eine unterschiedliche Anzahl von Befestigungsbolzen von der der massiven Teile der Zahnungen oder Wellungen der Enden der Wände der Kammer umfassen.
  8. Verbrennungskammer gemäß Anspruch 1 bis 7, dadurch gekennzeichnet, dass das ringförmige Gehäuse aus einem einzigen Teil oder aus zwei ringförmigen, jeweils radial internen oder externen Teilen realisiert ist.
  9. Turbomaschine, wie z. B. ein Turboreaktor oder ein Turboantrieb eines Flugzeugs, dadurch gekennzeichnet, dass sie eine ringförmige Verbrennungskammer gemäß einem der voranstehenden Ansprüche umfasst.
EP08075032.6A 2007-01-18 2008-01-14 Gasturbinenverbrennungskammer Active EP1956297B1 (de)

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
FR0700325A FR2911668B1 (fr) 2007-01-18 2007-01-18 Chambre de combustion d'une turbomachine

Publications (2)

Publication Number Publication Date
EP1956297A1 EP1956297A1 (de) 2008-08-13
EP1956297B1 true EP1956297B1 (de) 2016-03-30

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EP08075032.6A Active EP1956297B1 (de) 2007-01-18 2008-01-14 Gasturbinenverbrennungskammer

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US (1) US8087252B2 (de)
EP (1) EP1956297B1 (de)
CA (1) CA2619422C (de)
FR (1) FR2911668B1 (de)

Families Citing this family (5)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
FR2964725B1 (fr) * 2010-09-14 2012-10-12 Snecma Carenage aerodynamique pour fond de chambre de combustion
JP6266775B2 (ja) 2013-07-26 2018-01-24 エムアールエイ・システムズ・エルエルシー 航空機エンジンパイロン
FR3015639B1 (fr) * 2013-12-20 2018-08-31 Safran Aircraft Engines Chambre de combustion dans une turbomachine
GB201505502D0 (en) 2015-03-31 2015-05-13 Rolls Royce Plc Combustion equipment
DE102015224990A1 (de) 2015-12-11 2017-06-14 Rolls-Royce Deutschland Ltd & Co Kg Verfahren zur Montage einer Brennkammer eines Gasturbinentriebwerks

Family Cites Families (8)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US638290A (en) * 1899-08-14 1899-12-05 George J Hooper Rail-joint.
US5142871A (en) * 1991-01-22 1992-09-01 General Electric Company Combustor dome plate support having uniform thickness arcuate apex with circumferentially spaced coolant apertures
US5224825A (en) * 1991-12-26 1993-07-06 General Electric Company Locator pin retention device for floating joint
FR2686683B1 (fr) * 1992-01-28 1994-04-01 Snecma Turbomachine a chambre de combustion demontable.
US6904757B2 (en) * 2002-12-20 2005-06-14 General Electric Company Mounting assembly for the forward end of a ceramic matrix composite liner in a gas turbine engine combustor
FR2885201B1 (fr) * 2005-04-28 2010-09-17 Snecma Moteurs Chambre de combustion aisement demontable a performance aerodynamique amelioree
FR2887015B1 (fr) * 2005-06-14 2010-09-24 Snecma Moteurs Assemblage d'une chambre de combustion annulaire de turbomachine
FR2897144B1 (fr) * 2006-02-08 2008-05-02 Snecma Sa Chambre de combustion de turbomachine a fentes tangentielles

Also Published As

Publication number Publication date
CA2619422C (fr) 2015-11-17
FR2911668A1 (fr) 2008-07-25
CA2619422A1 (fr) 2008-07-18
EP1956297A1 (de) 2008-08-13
US8087252B2 (en) 2012-01-03
FR2911668B1 (fr) 2009-03-20
US20090293487A1 (en) 2009-12-03

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