EP1703207A1 - Brennkammerverkleidung für eine Turbomaschine - Google Patents

Brennkammerverkleidung für eine Turbomaschine Download PDF

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Publication number
EP1703207A1
EP1703207A1 EP06101397A EP06101397A EP1703207A1 EP 1703207 A1 EP1703207 A1 EP 1703207A1 EP 06101397 A EP06101397 A EP 06101397A EP 06101397 A EP06101397 A EP 06101397A EP 1703207 A1 EP1703207 A1 EP 1703207A1
Authority
EP
European Patent Office
Prior art keywords
fairing
flow
combustion chamber
central portion
flanks
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.)
Granted
Application number
EP06101397A
Other languages
English (en)
French (fr)
Other versions
EP1703207B1 (de
Inventor
Michel André Albert Desaulty
Michel Pierre Cazalens
Olivier Kreder
Alain Cayre
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 EP1703207A1 publication Critical patent/EP1703207A1/de
Application granted granted Critical
Publication of EP1703207B1 publication Critical patent/EP1703207B1/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
    • F23R3/02—Continuous combustion chambers using liquid or gaseous fuel characterised by the air-flow or gas-flow configuration
    • F23R3/04—Air inlet arrangements
    • F23R3/10—Air inlet arrangements for primary air
    • 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
    • F23R3/002—Wall structures

Definitions

  • the subject of this invention is a turbomachine combustion chamber shroud.
  • Such fairings cover the rear fuel injectors and protect them from shocks resulting from the ingestion of bodies such as ice blocks or birds in the machine. They have a substantially semi-toroidal shape and extend between two concentric edges for attachment to the edges of an annular bottom plate which borders the combustion chamber. The injectors extend through this plate. A central portion of the fairing is opened to allow the fuel injection hoses to pass to the injectors.
  • the openings may be a single circular slot (the fairing then being composed of two flanks, called “caps", concentric and separate) or consist of a succession of windows each leading to a group of injectors.
  • the combustion chamber in which the fairing extends often produces excessive noise associated with combustion instabilities and vibrations.
  • the reduction of acoustic emissions can be undertaken by adding stiffening or damping elements to the structure that produces them, to the detriment of the simplicity of manufacture, the lightness or the quality of the flow.
  • Other methods consist of dynamic steering of the combustion, but they do not yet have application in practice. Since it is difficult to obtain good results with these known methods, the restriction of the instabilities is sometimes neglected, which is however less and less acceptable because of the increasing demands of silence as well as good operation to which the engines must satisfy.
  • the fairings must also ensure a satisfactory flow of combustion air. Their rounded shape allows a smooth flow, provided with little turbulence, around them; but this favorable flow is guaranteed only at a nominal operating state outside which it is observed that the shape of the fairing is often no longer suitable: detachment of the flow can appear on certain portions of the sidewalls of the fairing, as well as inequalities of pressure.
  • the invention has been designed to overcome these shortcomings. It is based on improving the design of the fairing without adding material. Its essential feature is that at least one of the sides of the fairing is provided with at least one row of holes. The holes counteract the formation of a resonant cavity in the volume included in the fairing and reduce the noise emanating from it. According to another teaching of the invention, they also contribute, by stopping the pressure inequalities between the inside and the outside of the fairing, to regulate the flow of air for all modes of operation of the machine.
  • One aspect of the invention is a turbomachine combustion chamber shroud, covering a circular row of fuel injectors, provided with an open central portion and two flanks joining the central portion with two concentric fixing edges of the shroud.
  • an annular bottom plate of the combustion chamber characterized in that at least one of the sidewalls is provided with at least one row of bores.
  • a turbomachine combustion chamber comprising a housing, defining a diffusion chamber, a flame tube placed in the housing, a compressor diffuser opening into the diffusion chamber and constituting an origin of a first gas flow in the diffusion chamber, the flame tube comprising a ferrule and shroud joined to the ferrule and facing the compressor diffuser, the shroud covering a circular row of fuel injectors and being provided with a central portion and two concentric flanks joining the central portion to the ferrule, the first flow being directed from the diffuser to the central portion open, then bypassing the fairing along the flanks, and finally along the ferrule, characterized in that at least one flanks is provided with at least one row of bores.
  • Another aspect of the invention is a turbomachine equipped with this fairing or this combustion chamber.
  • FIG. 1 is a section along an axial plane of the machine, taken from one side only of the axis of rotation X of the rotor 1 of the machine.
  • This turbomachine is only partially represented in the fitted part of the invention, the remainder being not modified with respect to the known art.
  • a stator 3 of the machine Downstream of a high pressure compressor 2, a stator 3 of the machine comprises a diffuser 4 opening into a diffusion chamber 5 delimited by an outer casing 6, an internal casing 7 which is concentric with it and occupied by a flame tube.
  • the chamber bottom plate 11 carries fuel injectors 12 in connection with a fuel supply system 13 which supplies them through pipes 14 passing through the diffusion chamber 5 and the fuel injection system 13. 10. It is seen that the edges of the chamber bottom plate 11, the ferrule 9 and fairing 10 are assembled by bolts 15 superimposed in this order from the inside to the outside. The bolts 15 form two concentric circles and are associated with two edges of each of these parts.
  • the shroud 10 comprises two circular and concentric flanks 16 and 17 on either side of the openings through which the supply pipes 14 pass.
  • the flanks 16 and 17 are completely separated by an annular opening and are assembled separately to the rest of the stator.
  • the invention could equally well be applied to a one-piece fairing where the circular central slot would be replaced by a succession of shorter slots separated by radiating bridges joining the flanks 16 and 17 between them.
  • the flow of air at the outlet of the diffuser 4 preferably takes a path represented by the arrows and the flow lines of FIG. 2, which essentially by-passes the fairing 10 forming a flow which should be smooth on the along its flanks 16 and 17, that is to say, tangent to them over their entire length.
  • the flow of air from the diffuser 4 is directed first towards the center of the shroud 10. It bifurcates in front of the fairing 10 downstream of the turbomachine, then passes in front of the outer envelope and the inner envelope of the ferrule 9, which is thus refreshed. This main flow or first flow is completed by a second flow, also from the diffuser 4, which enters in the shroud 10 and the flame tube 8 through the central openings of the fairing 10.
  • Certain operating modes of the machine may however require a flow such as that of Figure 3, where a detachment 20 associated with a substantially air pocket stagnant occurs in front of a portion of the outer side of the outer flank 16 of the shroud 10. More generally, the detachment of the first flow often appears just downstream of a portion of greater curvature of the flanks 16 and 17 and especially the outer flank 16 no away from the connection to the ferrule 9.
  • the holes 21 may be circular or oblong, oval or rectangular, circular holes being easier to achieve. They are established in circular rows of the flanks 16 and 17 of the shroud 10, or only one of the flanks 16 or 17, with a regular distribution or not on the rows. A series of close circular holes gives a similar result to that of an oblong hole.
  • FIG. 4 represents a possible configuration of the invention, with a single row of bores 21. More complex patterns, associated with groups of bores, can give better results.
  • FIG. 5 represents a few of them, next to the elementary pattern (a) composed of a hole 21 unique of Figure 4, the patterns of two or three bores oriented axially (b or e), tangentially (c), triangle (d), square (f) or rhombus (g).
  • the rows of holes may comprise more or less regular combinations of this type of pattern.
  • An example is that of Figure 6, where patterns composed for example of eight bores close together and aligned in tangential direction alternate with triangles. Optimization depends on the concrete conditions of the flow and the importance of the improvement sought; it will be mostly determined empirically, so that there is no need to give rules beyond these examples.
  • the holes according to the invention must of course be distinguished from the holes in the edges of the fairing 10, which serve to receive the fixing bolts 15 to the bottom plate of chamber 11, so that they are plugged and do not have the properties of those of the invention; as well as holes established through the shell 9 of the flame tube 8, which are very numerous and of fine diameter, and whose role is to create an air flow under all circumstances to the flame tube 8 to maintain it at a moderate temperature while participating in the combustion when the home is reached.

Landscapes

  • Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Structures Of Non-Positive Displacement Pumps (AREA)
  • Fluidized-Bed Combustion And Resonant Combustion (AREA)
  • Portable Nailing Machines And Staplers (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
  • Combustion Methods Of Internal-Combustion Engines (AREA)
EP06101397A 2005-02-09 2006-02-08 Brennkammerverkleidung für eine Turbomaschine Expired - Lifetime EP1703207B1 (de)

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
FR0550379A FR2881813B1 (fr) 2005-02-09 2005-02-09 Carenage de chambre de combustion de turbomachine

Publications (2)

Publication Number Publication Date
EP1703207A1 true EP1703207A1 (de) 2006-09-20
EP1703207B1 EP1703207B1 (de) 2012-05-02

Family

ID=34953736

Family Applications (1)

Application Number Title Priority Date Filing Date
EP06101397A Expired - Lifetime EP1703207B1 (de) 2005-02-09 2006-02-08 Brennkammerverkleidung für eine Turbomaschine

Country Status (8)

Country Link
US (1) US7805943B2 (de)
EP (1) EP1703207B1 (de)
JP (1) JP2006220410A (de)
CN (1) CN1828141A (de)
CA (1) CA2535304C (de)
ES (1) ES2386150T3 (de)
FR (1) FR2881813B1 (de)
RU (1) RU2406932C2 (de)

Cited By (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
WO2013175126A2 (fr) 2012-05-25 2013-11-28 Snecma Virole de chambre de combustion de turbomachine
US10513984B2 (en) 2015-08-25 2019-12-24 General Electric Company System for suppressing acoustic noise within a gas turbine combustor

Families Citing this family (16)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
FR2943403B1 (fr) * 2009-03-17 2014-11-14 Snecma Chambre de combustion de turbomachine comprenant des moyens ameliores d'alimentation en air
US9291102B2 (en) 2011-09-07 2016-03-22 Siemens Energy, Inc. Interface ring for gas turbine fuel nozzle assemblies
FR2980554B1 (fr) * 2011-09-27 2013-09-27 Snecma Chambre annulaire de combustion d'une turbomachine
FR3019879A1 (fr) 2014-04-09 2015-10-16 Turbomeca Moteur d'aeronef comprenant un calage azimutal du diffuseur, par rapport a la chambre de combustion
DE102015206227A1 (de) * 2015-04-08 2016-10-13 Siemens Aktiengesellschaft Brenneranordnung
US10197275B2 (en) 2016-05-03 2019-02-05 General Electric Company High frequency acoustic damper for combustor liners
US10724739B2 (en) 2017-03-24 2020-07-28 General Electric Company Combustor acoustic damping structure
US10415480B2 (en) 2017-04-13 2019-09-17 General Electric Company Gas turbine engine fuel manifold damper and method of dynamics attenuation
US11149948B2 (en) 2017-08-21 2021-10-19 General Electric Company Fuel nozzle with angled main injection ports and radial main injection ports
US11156162B2 (en) 2018-05-23 2021-10-26 General Electric Company Fluid manifold damper for gas turbine engine
RU186956U1 (ru) * 2018-07-16 2019-02-11 Публичное Акционерное Общество "Одк-Сатурн" Жаровая труба камеры сгорания газотурбинного двигателя
FR3084141B1 (fr) * 2018-07-19 2021-04-02 Safran Aircraft Engines Ensemble pour une turbomachine
US11506125B2 (en) 2018-08-01 2022-11-22 General Electric Company Fluid manifold assembly for gas turbine engine
FR3095260B1 (fr) * 2019-04-18 2021-03-19 Safran Aircraft Engines Procede de definition de trous de passage d’air a travers une paroi de chambre de combustion
US12535212B2 (en) 2022-06-29 2026-01-27 General Electric Company Acoustic liner for a gas turbine engine
CN117917528A (zh) 2022-10-20 2024-04-23 通用电气公司 用于燃烧器的整流罩阻尼器

Citations (5)

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Publication number Priority date Publication date Assignee Title
US2901032A (en) * 1954-11-24 1959-08-25 Gen Thermique Procedes Brola S Combustion apparatus
EP0562792A1 (de) * 1992-03-23 1993-09-29 General Electric Company Schlagfester Aufsatz für eine Brennkammer
US5524430A (en) * 1992-01-28 1996-06-11 Societe National D'etude Et De Construction De Moteurs D'aviation S.N.E.C.M.A. Gas-turbine engine with detachable combustion chamber
DE19900025A1 (de) 1999-01-02 2000-07-06 Abb Research Ltd Brennerhaube
WO2004113794A1 (fr) * 2003-06-18 2004-12-29 Snecma Moteurs Chambre de combustion annulaire de turbomachine

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US2840989A (en) * 1955-09-15 1958-07-01 Gen Electric End cap for combustor
US3500639A (en) * 1968-09-10 1970-03-17 Gen Electric Combustion chamber mounting means
US3589127A (en) * 1969-02-04 1971-06-29 Gen Electric Combustion apparatus
GB2028488B (en) * 1978-08-19 1983-02-02 Rolls Royce Gas turbine
US5181379A (en) * 1990-11-15 1993-01-26 General Electric Company Gas turbine engine multi-hole film cooled combustor liner and method of manufacture
RU2103611C1 (ru) * 1995-07-11 1998-01-27 Открытое акционерное общество "Авиадвигатель" Кольцевая камера сгорания газотурбинного двигателя
RU2212589C1 (ru) * 2002-06-28 2003-09-20 Козырев Александр Валентинович Камера сгорания теплового двигателя
US6792757B2 (en) * 2002-11-05 2004-09-21 Honeywell International Inc. Gas turbine combustor heat shield impingement cooling baffle
US7062920B2 (en) * 2003-08-11 2006-06-20 General Electric Company Combustor dome assembly of a gas turbine engine having a free floating swirler

Patent Citations (5)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2901032A (en) * 1954-11-24 1959-08-25 Gen Thermique Procedes Brola S Combustion apparatus
US5524430A (en) * 1992-01-28 1996-06-11 Societe National D'etude Et De Construction De Moteurs D'aviation S.N.E.C.M.A. Gas-turbine engine with detachable combustion chamber
EP0562792A1 (de) * 1992-03-23 1993-09-29 General Electric Company Schlagfester Aufsatz für eine Brennkammer
DE19900025A1 (de) 1999-01-02 2000-07-06 Abb Research Ltd Brennerhaube
WO2004113794A1 (fr) * 2003-06-18 2004-12-29 Snecma Moteurs Chambre de combustion annulaire de turbomachine

Cited By (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
WO2013175126A2 (fr) 2012-05-25 2013-11-28 Snecma Virole de chambre de combustion de turbomachine
CN104520647A (zh) * 2012-05-25 2015-04-15 斯奈克玛 涡轮机燃烧室筒节
CN104520647B (zh) * 2012-05-25 2016-02-24 斯奈克玛 涡轮机燃烧室筒节
US10513984B2 (en) 2015-08-25 2019-12-24 General Electric Company System for suppressing acoustic noise within a gas turbine combustor

Also Published As

Publication number Publication date
FR2881813A1 (fr) 2006-08-11
RU2006103679A (ru) 2007-08-20
RU2406932C2 (ru) 2010-12-20
US7805943B2 (en) 2010-10-05
EP1703207B1 (de) 2012-05-02
FR2881813B1 (fr) 2011-04-08
US20060174626A1 (en) 2006-08-10
CA2535304C (fr) 2015-03-31
ES2386150T3 (es) 2012-08-10
JP2006220410A (ja) 2006-08-24
CN1828141A (zh) 2006-09-06
CA2535304A1 (fr) 2006-08-09

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