EP1777458A1 - Leistungsverbesserung einer Brennkammer durch vielfache Perforierung der Wände - Google Patents

Leistungsverbesserung einer Brennkammer durch vielfache Perforierung der Wände Download PDF

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Publication number
EP1777458A1
EP1777458A1 EP06120816A EP06120816A EP1777458A1 EP 1777458 A1 EP1777458 A1 EP 1777458A1 EP 06120816 A EP06120816 A EP 06120816A EP 06120816 A EP06120816 A EP 06120816A EP 1777458 A1 EP1777458 A1 EP 1777458A1
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EP
European Patent Office
Prior art keywords
holes
wall
cooling
combustion chamber
bores
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
EP06120816A
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English (en)
French (fr)
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EP1777458B1 (de
Inventor
Daniel Bernier
Jean-Michel Campion
Stéphane Touchaud
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 date
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Publication of EP1777458A1 publication Critical patent/EP1777458A1/de
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Publication of EP1777458B1 publication Critical patent/EP1777458B1/de
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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
    • F23R3/002—Wall structures
    • 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
    • F23R3/06—Arrangement of apertures along the flame tube
    • 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
    • F23R2900/03041—Effusion cooled combustion chamber walls or domes
    • 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
    • F23R2900/03042—Film cooled combustion chamber walls or domes

Definitions

  • the present invention relates to the general field of turbomachine combustion chambers. It relates more particularly to an annular wall for combustion chamber cooled by a so-called "multiperforation" process.
  • annular turbomachine combustion chamber is formed of an inner annular wall and an outer annular wall which are connected upstream by a transverse wall forming a chamber bottom.
  • the inner and outer walls are each provided with a plurality of holes and various orifices allowing air circulating around the combustion chamber to penetrate inside thereof.
  • so-called “primary” and “dilution” holes are formed in these walls to convey air inside the combustion chamber.
  • the air passing through the primary holes helps to create an air / fuel mixture that is burned in the chamber, while the air from the dilution holes is intended to promote the dilution of the same air / fuel mixture.
  • the inner and outer walls which are generally metallic, are subject to the high temperatures of the gases from the combustion of the air / fuel mixture.
  • additional holes called multiperforation holes are also drilled through the walls over their entire surface. These multiperforation holes allow the air circulating outside the chamber to penetrate inside thereof by forming along the walls of the cooling air films.
  • the main object of the present invention is therefore to overcome such drawbacks by proposing an annular wall of a combustion chamber provided with additional bores intended to cool the zones situated directly downstream of the primary and dilution holes.
  • annular wall of a turbomachine combustion chamber having a cold side and a hot side, the wall being provided with a plurality of primary holes and dilution holes to allow circulating air the cold side of the wall to enter the hot side to respectively provide combustion and dilution of an air / fuel mixture, the primary holes and the dilution holes being distributed in circumferential rows, and a plurality of cooling orifices for allowing the air flowing from the cold side of the wall to enter the warm side to form a cooling air film along said wall, the cooling orifices being distributed in a plurality of circumferential rows; axially spaced from each other, the number of cooling orifices being identical in each row, characterized in that it further comprises a plurality of of bores arranged directly downstream of the primary holes and the dilution holes and distributed in circumferential rows, the bores of the same row having a substantially identical diameter, being spaced at a constant pitch and having intrinsic characteristics different from those of the cooling holes of adjacent
  • intrinsic characteristics of the holes we mean the number, the inclination and the diameter of these holes.
  • the presence of holes having intrinsic characteristics different from those of the cooling orifices and arranged directly downstream of the primary and dilution holes makes it possible to ensure efficient cooling of these zones. Any risk of formation of cracks is thus avoided.
  • the specific holes are distributed in circumferential rows, have the same diameter and are spaced a constant pitch which greatly facilitates the drilling operations and thus reduces the costs and manufacturing time of the wall.
  • the number of holes in the same row may be different from the number of cooling orifices of the adjacent rows.
  • the inclination of the bores of the same row relative to a normal to the wall may be different from that of the cooling orifices of adjacent rows.
  • the diameter of the bores of the same row may be different from that of the cooling holes of adjacent rows.
  • the present invention also relates to a combustion chamber and a turbomachine (having a combustion chamber) comprising an annular wall as defined above.
  • FIG. 1 illustrates a combustion chamber for a turbomachine.
  • a turbomachine comprises in particular a compression section (not shown) in which air is compressed before being injected into a chamber housing 2, then into a combustion chamber 4 mounted inside thereof.
  • Compressed air is introduced into the combustion chamber and mixed with fuel before being burned.
  • the gases resulting from this combustion are then directed to a high-pressure turbine 5 disposed at the outlet of the combustion chamber 4.
  • the combustion chamber 4 is of annular type. It is formed of an inner annular wall 6 and an outer annular wall 8 which are connected upstream by a transverse wall 10 forming a chamber bottom.
  • the inner 6 and outer 8 walls extend along a longitudinal axis X-X slightly inclined relative to the longitudinal axis Y-Y of the turbomachine.
  • the chamber bottom 10 is provided with a plurality of openings 12 in which fuel injectors 14 are mounted.
  • the chamber casing 2 which is formed of an inner casing 2a and an outer casing 2b, furnishes with the combustion chamber 4 an annular space 16 in which compressed air for combustion is admitted to the chamber. dilution and cooling of the chamber.
  • the inner 6 and outer 8 walls each have a cold side 6a, 8a disposed on the side of the annular space 16 in which the compressed air circulates and a hot side 6b, 8b facing the inside of the combustion chamber 4 ( Figure 3).
  • the combustion chamber 4 is divided into a so-called “primary” zone (or combustion zone) and a so-called “secondary” zone (or dilution zone) located downstream of the previous one (the downstream means with respect to the general direction of flow of gases from the combustion of the air / fuel mixture inside the combustion chamber).
  • the air which supplies the primary zone of the combustion chamber 4 is introduced by one or more circumferential rows of primary holes 18 made in the inner and outer walls 6 of the chamber.
  • the air supplying the secondary zone of the chamber it borrows a plurality of dilution holes 20 also formed in the inner 6 and outer 8 walls. These dilution holes 20 are aligned in one or more circumferential rows which are axially offset. downstream from the rows of primary holes 18.
  • the primary holes 18 and the dilution holes 20 are distributed on the inner 6 and outer 8 walls in rows extending over the entire circumference of the walls.
  • a plurality of cooling orifices 22 are provided (FIGS. 2 and 3).
  • These orifices 22, which provide a cooling of the walls 6, 8 by multiperforation, are distributed in a plurality of circumferential rows spaced axially from each other. These rows of multiperforation holes cover almost the entire surface of the walls 6, 8 of the chamber.
  • the number and the diameter d1 of the cooling orifices 22 are identical in each of the rows.
  • the pitch p1 between two orifices 22 of the same row is also identical throughout the row.
  • the adjacent rows of cooling orifices are arranged so that the orifices 22 are arranged in staggered rows as shown in FIG.
  • the cooling orifices 22 generally have an angle of inclination ⁇ with respect to a normal N to the annular wall 6, 8 through which they are pierced.
  • This inclination ⁇ allows the air passing through these orifices to form a film of air along the hot side 6b, 8b of the annular wall 6, 8.
  • it also makes it possible to increase the surface area. of the annular wall which is cooled.
  • the inclination ⁇ of the cooling orifices 22 is directed so that the film of air thus formed flows in the direction of flow of the combustion gases inside the chamber (shown schematically by the arrow on Figure 3).
  • the diameter d 1 of the cooling orifices 22 may be between 0.3 and 1. mm, the pitch p1 between 1 and 10 mm and their inclination ⁇ between -80 ° and + 80 °.
  • primary holes 18 and the dilution holes 20 have a diameter of the order of 5 to 20 mm.
  • each annular wall 6, 8 of the combustion chamber further comprises a plurality of bores 24 which are arranged directly downstream of the primary holes 18 and dilution 20 and which are distributed in circumferential rows.
  • the bores 24 of the same row have a substantially identical diameter d2 , are spaced at a constant pitch P2 and have intrinsic characteristics different from those of the cooling orifices 22 of adjacent rows.
  • these holes 24 are thus distributed in one or more rows (for example from 1 to 3 rows) which are arranged directly downstream of said hole 18, 20.
  • the intrinsic characteristics of these holes 24 are different from those of the cooling orifices 22, that is to say that the number of holes in the same row is different from that of a row of cooling orifices, and / or or the inclination of the bores of the same row with respect to a normal N to the wall 6, 8 is different from that of the cooling orifices, and / or the diameter d2 of the bores of the same row is different from that of d1 22. It should be noted that these three intrinsic characteristics of the holes 24 can be added.
  • the number of bores 24 on the same row may be, over the entire circumference of the wall, of the order of 860 when the number of cooling orifices 22 is of the order of 576. .
  • the inclination of the bores 24 relative to a normal to the walls 6, 8 is zero (that is to say that the bores are substantially perpendicular to the walls), while the inclination ⁇ of the cooling orifices 22 with respect to this same normal is between 30 ° and 70 °.
  • the bores 24 of the same row have an identical diameter d2 and are spaced by a constant pitch p2 .
  • Such holes are typically made by laser using a machine programmed according to the position of each of the holes to be made.
  • the characteristics of the holes according to the invention make it possible, with respect to a localized treatment (for which the bores are made only in the direct vicinity of each of the primary and dilution holes), to considerably simplify the programming of the machine, and thus to reduce the costs and manufacturing delays.

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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)
EP06120816.1A 2005-10-18 2006-09-18 Leistungsverbesserung einer Brennkammer durch vielfache Perforierung der Wände Active EP1777458B1 (de)

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
FR0510584A FR2892180B1 (fr) 2005-10-18 2005-10-18 Amelioration des perfomances d'une chambre de combustion par multiperforation des parois

Publications (2)

Publication Number Publication Date
EP1777458A1 true EP1777458A1 (de) 2007-04-25
EP1777458B1 EP1777458B1 (de) 2015-08-12

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Country Status (4)

Country Link
US (1) US7748222B2 (de)
EP (1) EP1777458B1 (de)
FR (1) FR2892180B1 (de)
RU (1) RU2413134C2 (de)

Cited By (7)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
FR2922629A1 (fr) * 2007-10-22 2009-04-24 Snecma Sa Chambre de combustion a dilution optimisee et turbomachine en etant munie
FR2974162A1 (fr) * 2011-04-14 2012-10-19 Snecma Virole de tube a flamme dans une chambre de combustion de turbomachine
FR2975465A1 (fr) * 2011-05-19 2012-11-23 Snecma Paroi pour chambre de combustion de turbomachine comprenant un agencement optimise d'orifices d'entree d'air
RU2478875C2 (ru) * 2007-10-22 2013-04-10 Снекма Стенка камеры сгорания с оптимизированным разжижением и охлаждением, камера сгорания и газотурбинный двигатель, снабженный такой стенкой
FR2981733A1 (fr) * 2011-10-25 2013-04-26 Snecma Module de chambre de combustion de turbomachine d'aeronef et procede de conception de celui-ci
RU2584746C2 (ru) * 2011-02-25 2016-05-20 Снекма Кольцевая камера сгорания для газотурбинного двигателя, содержащая улучшенные отверстия для охлаждения
EP3040615A1 (de) * 2014-12-17 2016-07-06 United Technologies Corporation Vorrichtung und system zur regelung der aktiven brennkammerverdünnungslochwärmeübertragung

Families Citing this family (27)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US7614235B2 (en) * 2005-03-01 2009-11-10 United Technologies Corporation Combustor cooling hole pattern
US8171634B2 (en) * 2007-07-09 2012-05-08 Pratt & Whitney Canada Corp. Method of producing effusion holes
US7905094B2 (en) * 2007-09-28 2011-03-15 Honeywell International Inc. Combustor systems with liners having improved cooling hole patterns
US8616004B2 (en) * 2007-11-29 2013-12-31 Honeywell International Inc. Quench jet arrangement for annular rich-quench-lean gas turbine combustors
US8127554B2 (en) * 2007-11-29 2012-03-06 Honeywell International Inc. Quench jet arrangement for annular rich-quench-lean gas turbine combustors
US8056342B2 (en) * 2008-06-12 2011-11-15 United Technologies Corporation Hole pattern for gas turbine combustor
US8091367B2 (en) * 2008-09-26 2012-01-10 Pratt & Whitney Canada Corp. Combustor with improved cooling holes arrangement
US8438856B2 (en) * 2009-03-02 2013-05-14 General Electric Company Effusion cooled one-piece can combustor
US9062884B2 (en) 2011-05-26 2015-06-23 Honeywell International Inc. Combustors with quench inserts
FR2982008B1 (fr) * 2011-10-26 2013-12-13 Snecma Paroi annulaire de chambre de combustion a refroidissement ameliore au niveau des trous primaires et de dilution
FR2991028B1 (fr) * 2012-05-25 2014-07-04 Snecma Virole de chambre de combustion de turbomachine
US10260748B2 (en) * 2012-12-21 2019-04-16 United Technologies Corporation Gas turbine engine combustor with tailored temperature profile
EP2971966B1 (de) * 2013-03-15 2017-04-19 Rolls-Royce Corporation Innenverkleidung für eine gasturbinenbrennkammer
WO2014149081A1 (en) * 2013-03-15 2014-09-25 Rolls-Royce Corporation Counter swirl doublet combustor
WO2015126501A2 (en) 2013-12-06 2015-08-27 United Technologies Corporation Co-swirl orientation of combustor effusion passages for gas turbine engine combustor
FR3035707B1 (fr) * 2015-04-29 2019-11-01 Safran Aircraft Engines Chambre de combustion coudee d'une turbomachine
US10670267B2 (en) * 2015-08-14 2020-06-02 Raytheon Technologies Corporation Combustor hole arrangement for gas turbine engine
GB201518345D0 (en) * 2015-10-16 2015-12-02 Rolls Royce Combustor for a gas turbine engine
JP6026028B1 (ja) * 2016-03-10 2016-11-16 三菱日立パワーシステムズ株式会社 燃焼器用パネル、燃焼器、燃焼装置、ガスタービン、及び燃焼器用パネルの冷却方法
DE102016219424A1 (de) * 2016-10-06 2018-04-12 Rolls-Royce Deutschland Ltd & Co Kg Brennkammeranordnung einer Gasturbine sowie Fluggasturbine
US11015529B2 (en) 2016-12-23 2021-05-25 General Electric Company Feature based cooling using in wall contoured cooling passage
US20180266687A1 (en) * 2017-03-16 2018-09-20 General Electric Company Reducing film scrubbing in a combustor
US10816202B2 (en) 2017-11-28 2020-10-27 General Electric Company Combustor liner for a gas turbine engine and an associated method thereof
US11255543B2 (en) 2018-08-07 2022-02-22 General Electric Company Dilution structure for gas turbine engine combustor
DE102019105442A1 (de) * 2019-03-04 2020-09-10 Rolls-Royce Deutschland Ltd & Co Kg Verfahren zur Herstellung eines Triebwerksbauteils mit einer Kühlkanalanordnung und Triebwerksbauteil
EP3848556A1 (de) * 2020-01-13 2021-07-14 Ansaldo Energia Switzerland AG Gasturbinentriebwerk mit übergangsstück mit geneigten kühlöffnungen
CN116202106B (zh) * 2023-03-08 2024-05-03 中国科学院工程热物理研究所 一种气膜孔与掺混孔耦合设计的发动机燃烧室火焰筒结构

Citations (6)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
FR1493144A (fr) * 1966-08-19 1967-08-25 Lucas Industries Ltd Perfectionnements aux appareils de combustion pour moteurs de turbines à gaz
US5261223A (en) * 1992-10-07 1993-11-16 General Electric Company Multi-hole film cooled combustor liner with rectangular film restarting holes
US5775108A (en) * 1995-04-26 1998-07-07 Societe Nationale D'etude Et De Construction De Moteurs D'aviation "Snecma" Combustion chamber having a multi-hole cooling system with variably oriented holes
EP0943868A2 (de) * 1998-03-18 1999-09-22 General Electric Company Gasturbinenbrennkammer
EP1001222A2 (de) * 1998-11-13 2000-05-17 General Electric Company Brennkammerhemd mit Filmkühlung
US6145319A (en) * 1998-07-16 2000-11-14 General Electric Company Transitional multihole combustion liner

Family Cites Families (17)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3572031A (en) * 1969-07-11 1971-03-23 United Aircraft Corp Variable area cooling passages for gas turbine burners
GB1492049A (en) * 1974-12-07 1977-11-16 Rolls Royce Combustion equipment for gas turbine engines
US4244178A (en) * 1978-03-20 1981-01-13 General Motors Corporation Porous laminated combustor structure
US4180972A (en) * 1978-06-08 1980-01-01 General Motors Corporation Combustor support structure
US4269032A (en) * 1979-06-13 1981-05-26 General Motors Corporation Waffle pattern porous material
US4302940A (en) * 1979-06-13 1981-12-01 General Motors Corporation Patterned porous laminated material
US4296606A (en) * 1979-10-17 1981-10-27 General Motors Corporation Porous laminated material
US5241827A (en) * 1991-05-03 1993-09-07 General Electric Company Multi-hole film cooled combuster linear with differential cooling
US6474070B1 (en) * 1998-06-10 2002-11-05 General Electric Company Rich double dome combustor
US6434821B1 (en) * 1999-12-06 2002-08-20 General Electric Company Method of making a combustion chamber liner
US6427446B1 (en) * 2000-09-19 2002-08-06 Power Systems Mfg., Llc Low NOx emission combustion liner with circumferentially angled film cooling holes
US6408629B1 (en) * 2000-10-03 2002-06-25 General Electric Company Combustor liner having preferentially angled cooling holes
US6620457B2 (en) * 2001-07-13 2003-09-16 General Electric Company Method for thermal barrier coating and a liner made using said method
US6513331B1 (en) * 2001-08-21 2003-02-04 General Electric Company Preferential multihole combustor liner
US7093439B2 (en) * 2002-05-16 2006-08-22 United Technologies Corporation Heat shield panels for use in a combustor for a gas turbine engine
US7216485B2 (en) * 2004-09-03 2007-05-15 General Electric Company Adjusting airflow in turbine component by depositing overlay metallic coating
US7895841B2 (en) * 2006-07-14 2011-03-01 General Electric Company Method and apparatus to facilitate reducing NOx emissions in turbine engines

Patent Citations (6)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
FR1493144A (fr) * 1966-08-19 1967-08-25 Lucas Industries Ltd Perfectionnements aux appareils de combustion pour moteurs de turbines à gaz
US5261223A (en) * 1992-10-07 1993-11-16 General Electric Company Multi-hole film cooled combustor liner with rectangular film restarting holes
US5775108A (en) * 1995-04-26 1998-07-07 Societe Nationale D'etude Et De Construction De Moteurs D'aviation "Snecma" Combustion chamber having a multi-hole cooling system with variably oriented holes
EP0943868A2 (de) * 1998-03-18 1999-09-22 General Electric Company Gasturbinenbrennkammer
US6145319A (en) * 1998-07-16 2000-11-14 General Electric Company Transitional multihole combustion liner
EP1001222A2 (de) * 1998-11-13 2000-05-17 General Electric Company Brennkammerhemd mit Filmkühlung

Cited By (13)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
RU2478875C2 (ru) * 2007-10-22 2013-04-10 Снекма Стенка камеры сгорания с оптимизированным разжижением и охлаждением, камера сгорания и газотурбинный двигатель, снабженный такой стенкой
FR2922629A1 (fr) * 2007-10-22 2009-04-24 Snecma Sa Chambre de combustion a dilution optimisee et turbomachine en etant munie
EP2053312A3 (de) * 2007-10-22 2014-10-29 Snecma Brennkammer mit optimierter Verdünnung und damit ausgestattetes Turbotriebwerk
RU2474763C2 (ru) * 2007-10-22 2013-02-10 Снекма Камера сгорания с оптимизированным разбавлением и турбомашина, снабженная такой камерой сгорания
RU2584746C2 (ru) * 2011-02-25 2016-05-20 Снекма Кольцевая камера сгорания для газотурбинного двигателя, содержащая улучшенные отверстия для охлаждения
FR2974162A1 (fr) * 2011-04-14 2012-10-19 Snecma Virole de tube a flamme dans une chambre de combustion de turbomachine
FR2975465A1 (fr) * 2011-05-19 2012-11-23 Snecma Paroi pour chambre de combustion de turbomachine comprenant un agencement optimise d'orifices d'entree d'air
FR2981733A1 (fr) * 2011-10-25 2013-04-26 Snecma Module de chambre de combustion de turbomachine d'aeronef et procede de conception de celui-ci
WO2013060985A1 (fr) * 2011-10-25 2013-05-02 Snecma Module de chambre de combustion de turbomachine d'aéronef et procédé de conception de celui-ci
GB2511440A (en) * 2011-10-25 2014-09-03 Snecma Aircraft turbomachine combustion chamber module and method for designing same
US9765970B2 (en) 2011-10-25 2017-09-19 Safran Aircraft Engines Aircraft turbomachine combustion chamber module and method for designing same
GB2511440B (en) * 2011-10-25 2019-02-27 Snecma Aircraft turbomachine combustion chamber module having improved air distribution within the combustion chamber and method for designing same
EP3040615A1 (de) * 2014-12-17 2016-07-06 United Technologies Corporation Vorrichtung und system zur regelung der aktiven brennkammerverdünnungslochwärmeübertragung

Also Published As

Publication number Publication date
US20070084219A1 (en) 2007-04-19
FR2892180A1 (fr) 2007-04-20
EP1777458B1 (de) 2015-08-12
RU2413134C2 (ru) 2011-02-27
US7748222B2 (en) 2010-07-06
RU2006136873A (ru) 2008-04-27
FR2892180B1 (fr) 2008-02-01

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