EP1777458B1 - 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
EP1777458B1
EP1777458B1 EP06120816.1A EP06120816A EP1777458B1 EP 1777458 B1 EP1777458 B1 EP 1777458B1 EP 06120816 A EP06120816 A EP 06120816A EP 1777458 B1 EP1777458 B1 EP 1777458B1
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EP
European Patent Office
Prior art keywords
holes
wall
combustion chamber
bores
cooling
Prior art date
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Active
Application number
EP06120816.1A
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English (en)
French (fr)
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EP1777458A1 (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 of EP1777458A1 publication Critical patent/EP1777458A1/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.
  • annular wall of turbomachine combustion chamber being provided with a plurality of primary holes and dilution holes. Cooling orifices are distributed in circumferential rows, the number of these orifices being identical in each row. A plurality of bores are disposed directly downstream of the primary holes and the dilution holes. These holes are distributed in circumferential rows. The bores of the same row have a substantially identical diameter, and are spaced a constant pitch throughout the row. Variations in inclination or frequency of said holes relative to the other cooling orifices are not discussed in the document.
  • Patent documents EP 1001222 and US 5775108 describe similar annular combustion chamber walls, and there are discussed variations in number (EP) or inclination (US) of the cooling orifices.
  • 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 according to claim 1.
  • 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 effective cooling of these areas. Any risk of formation of cracks is thus avoided.
  • the specific bores are distributed in circumferential rows, have the same diameter and are spaced a constant pitch which greatly facilitates 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, and / or the inclination of the bores of the same row relative to a normal to the wall may be different from that of the cooling holes in adjacent rows.
  • the diameter of the bores of the same row may be different from that of the cooling holes of the 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.
  • the figure 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 ( Figures 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. figure 1 .
  • 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 the 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 holes are substantially perpendicular to the walls), while the inclination ⁇ of the cooling orifices 22 relative to at 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)

Claims (3)

  1. Ringförmige Wand (6, 8) einer Brennkammer (4) einer Turbomaschine, umfassend eine kalte Seite (6a, 8a) und eine heiße Seite (6b, 8b), wobei die Wand versehen ist:
    mit einer Vielzahl von Primärlöchern (18) und von Verdünnungslöchern (20), um Luft, die auf der kalten Seite (6a, 8a) der Wand (6, 8) zirkuliert, zu ermöglichen, auf die heiße Seite (6b, 8b) zu dringen, um die Verbrennung bzw. die Verdünnung eines Luft/Treibstoff-Gemischs sicherzustellen, wobei die Primärlöcher (18) und die Verdünnungslöcher (20) entlang von Umfangsreihen verteilt sind,
    mit einer Vielzahl von Kühlöffnungen (22), um der auf der kalten Seite (6a, 8a) der Wand (6, 8) zirkulierenden Luft zu ermöglichen, auf die heiße Seite (6b, 8b) zu dringen, um einen Kühlluftfilm entlang der Wand auszubilden, wobei die Kühlöffnungen (22) entlang einer Vielzahl von Umfangsreihen, die axial voneinander beabstandet sind, verteilt sind, wobei die Anzahl an Kühlöffnungen (22) in jeder Reihe identisch ist, und
    mit einer Vielzahl von Bohrungen (24), die direkt stromabwärts der Primärlöcher (18) und der Verdünnungslöcher (20) angeordnet und entlang von Umfangsreihen verteilt sind, wobei die Bohrungen (24) einer gleichen Reihe einen im Wesentlichen identischen Durchmesser (d2) aufweisen und um einen kostanten Abstand (p2) beabstandet sind,
    wobei die Anzahl an Bohrungen (24) einer gleichen Reihe von der Anzahl an Kühlöffnungen (22) der benachbarten Reihen verschieden ist, und/oder
    wobei die Neigung der Bohrungen (24) einer gleichen Reihe gegenüber einer Normalen (N) zu der Wand von derjenigen der Kühlöffnungen (22) der benachbarten Reihen verschieden ist.
  2. Brennkammer (4) einer Turbomaschine, umfassend wenigstens eine ringförmige Wand (6, 8) nach Anspruch 1.
  3. Turbomaschine, umfassend eine Brennkammer (4) mit wenigstens einer ringförmigen Wand (6, 8) nach Anspruch 1.
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 EP1777458A1 (de) 2007-04-25
EP1777458B1 true EP1777458B1 (de) 2015-08-12

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Country Status (4)

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

Families Citing this family (34)

* 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
FR2922630B1 (fr) * 2007-10-22 2015-11-13 Snecma Paroi de chambre de combustion a dilution et refroidissement optimises,chambre de combustion et turbomachine en etant munies
FR2922629B1 (fr) * 2007-10-22 2009-12-25 Snecma Chambre de combustion a dilution optimisee et turbomachine en etant munie
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
FR2972027B1 (fr) * 2011-02-25 2013-03-29 Snecma Chambre annulaire de combustion de turbomachine comprenant des orifices de dilution ameliores
FR2974162B1 (fr) * 2011-04-14 2018-04-13 Safran Aircraft Engines Virole de tube a flamme dans une chambre de combustion de turbomachine
FR2975465B1 (fr) * 2011-05-19 2018-03-09 Safran Aircraft Engines Paroi pour chambre de combustion de turbomachine comprenant un agencement optimise d'orifices d'entree d'air
US9062884B2 (en) 2011-05-26 2015-06-23 Honeywell International Inc. Combustors with quench inserts
FR2981733B1 (fr) 2011-10-25 2013-12-27 Snecma Module de chambre de combustion de turbomachine d'aeronef et procede de conception de celui-ci
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
WO2014149081A1 (en) * 2013-03-15 2014-09-25 Rolls-Royce Corporation Counter swirl doublet combustor
EP2971966B1 (de) * 2013-03-15 2017-04-19 Rolls-Royce Corporation Innenverkleidung für eine gasturbinenbrennkammer
EP3077728B8 (de) 2013-12-06 2021-03-31 Raytheon Technologies Corporation Gasturbinenbrennkammer mit co-swirl-ausrichtung von effusionslöchern, und methode
US20160178199A1 (en) * 2014-12-17 2016-06-23 United Technologies Corporation Combustor dilution hole active heat transfer control apparatus and system
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 (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP1632720A1 (de) * 2004-09-03 2006-03-08 General Electric Company Einstellen der Luftströmungsmenge an einem Bauteil einer Turbine durch Aufbringen einer metallischen Beschichtung

Family Cites Families (25)

* 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
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
US4302940A (en) * 1979-06-13 1981-12-01 General Motors Corporation Patterned porous laminated material
US4269032A (en) * 1979-06-13 1981-05-26 General Motors Corporation Waffle pattern porous material
US4296606A (en) * 1979-10-17 1981-10-27 General Motors Corporation Porous laminated material
SU1373045A1 (ru) * 1986-05-26 1996-12-20 В.М. Кофман Охлаждаемый корпус
US5181379A (en) * 1990-11-15 1993-01-26 General Electric Company Gas turbine engine multi-hole film cooled combustor liner and method of manufacture
US5241827A (en) * 1991-05-03 1993-09-07 General Electric Company Multi-hole film cooled combuster linear with differential cooling
US5261223A (en) * 1992-10-07 1993-11-16 General Electric Company Multi-hole film cooled combustor liner with rectangular film restarting holes
FR2733582B1 (fr) * 1995-04-26 1997-06-06 Snecma Chambre de combustion comportant une multiperforation d'inclinaison axiale et tangentielle variable
US6192689B1 (en) * 1998-03-18 2001-02-27 General Electric Company Reduced emissions gas turbine combustor
US6474070B1 (en) * 1998-06-10 2002-11-05 General Electric Company Rich double dome combustor
US6145319A (en) * 1998-07-16 2000-11-14 General Electric Company Transitional multihole combustion liner
US6205789B1 (en) * 1998-11-13 2001-03-27 General Electric Company Multi-hole film cooled combuster liner
RU2162194C1 (ru) * 1999-11-24 2001-01-20 Общество с ограниченной ответственностью Научно-производственное предприятие "ЭСТ" Камера сгорания
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
US7895841B2 (en) * 2006-07-14 2011-03-01 General Electric Company Method and apparatus to facilitate reducing NOx emissions in turbine engines

Patent Citations (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP1632720A1 (de) * 2004-09-03 2006-03-08 General Electric Company Einstellen der Luftströmungsmenge an einem Bauteil einer Turbine durch Aufbringen einer metallischen Beschichtung

Also Published As

Publication number Publication date
US20070084219A1 (en) 2007-04-19
FR2892180B1 (fr) 2008-02-01
EP1777458A1 (de) 2007-04-25
RU2413134C2 (ru) 2011-02-27
FR2892180A1 (fr) 2007-04-20
RU2006136873A (ru) 2008-04-27
US7748222B2 (en) 2010-07-06

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