EP0743490A1 - Brennkammer mit einer Vielzahl von Filmkühlungsbohrungen die in verschiedene axiale und tangentiale Richtungen geneigt sind - Google Patents

Brennkammer mit einer Vielzahl von Filmkühlungsbohrungen die in verschiedene axiale und tangentiale Richtungen geneigt sind Download PDF

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Publication number
EP0743490A1
EP0743490A1 EP96400863A EP96400863A EP0743490A1 EP 0743490 A1 EP0743490 A1 EP 0743490A1 EP 96400863 A EP96400863 A EP 96400863A EP 96400863 A EP96400863 A EP 96400863A EP 0743490 A1 EP0743490 A1 EP 0743490A1
Authority
EP
European Patent Office
Prior art keywords
orifices
zones
wall
flow
axial
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
EP96400863A
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English (en)
French (fr)
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EP0743490B1 (de
Inventor
Denis Roger Henri Ansart
Patrick Samuel André Ciccia
Michel André Albert Desaulty
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
Societe Nationale dEtude et de Construction de Moteurs dAviation SNECMA
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 Societe Nationale dEtude et de Construction de Moteurs dAviation SNECMA, SNECMA SAS filed Critical Societe Nationale dEtude et de Construction de Moteurs dAviation SNECMA
Publication of EP0743490A1 publication Critical patent/EP0743490A1/de
Application granted granted Critical
Publication of EP0743490B1 publication Critical patent/EP0743490B1/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
    • F23R3/06—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
    • F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
    • F23R3/002—Wall structures
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05B—INDEXING SCHEME RELATING TO WIND, SPRING, WEIGHT, INERTIA OR LIKE MOTORS, TO MACHINES OR ENGINES FOR LIQUIDS COVERED BY SUBCLASSES F03B, F03D AND F03G
    • F05B2250/00—Geometry
    • F05B2250/30—Arrangement of components
    • F05B2250/32—Arrangement of components according to their shape
    • F05B2250/322—Arrangement of components according to their shape tangential
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05B—INDEXING SCHEME RELATING TO WIND, SPRING, WEIGHT, INERTIA OR LIKE MOTORS, TO MACHINES OR ENGINES FOR LIQUIDS COVERED BY SUBCLASSES F03B, F03D AND F03G
    • F05B2260/00—Function
    • F05B2260/20—Heat transfer, e.g. cooling
    • F05B2260/202—Heat transfer, e.g. cooling by film cooling

Definitions

  • the present invention relates to a combustion chamber, in particular of a turbomachine, which is delimited by at least one axial wall provided with a plurality of through orifices constituting a "multi-perforation" intended, in particular, for the passage of a refrigeration fluid from said axial wall, and provided with a plurality of dilution holes regularly distributed in a transverse plane with respect to the general direction of the flow of the burnt gases originating from the combustion, each orifice having a geometric axis inclined at an angle A by with respect to the normal to said wall, said genometric axis being arranged in a plane containing said normal which makes an angle B with respect to the plane defined by said normal and the general direction of flow of the burnt gases.
  • the cooling mode by multiperforation is known.
  • the orifices are generally arranged in staggered rows with an equidistant network of meshes.
  • These ports are supplied with cooling air from the compressor.
  • the heat exchanges involved are then, forced convection inside the orifices, conduction within the wall itself.
  • the supply of cooling air to these orifices generates, downstream of the flow, on the internal part of the wall, a protective film between the chamber wall and the burnt gases from the combustion. In order to limit the degradation of the efficiency of this film, it is ensured that the cooling air does not mix too soon with the burnt gases.
  • the orifices are inclined at an angle A relative to the normal to the internal wall, so that the cooling air comes to lick this wall to be cooled.
  • EP-A-0 486 133 discloses a wall of this type, in which the orifices are inclined in axial planes.
  • EP-A-0 492 864 further reveals that the orifices are also inclined at a tangential angle B which generally coincides with the angle of the vortex of the combustion gases along the internal surface of the wall.
  • EP-A-0 592 161 shows in FIG. 6 a multi-perforated annular wall of a combustion chamber in which the orifices are defined by an axial inclination A and a tangential angle B of in such a way that the flow of fresh air introduced into the chamber creates a ring of protective air which swirls around the flow of the burnt gases.
  • 3D calculations show that the flow of gases in the combustion chamber is not always longitudinal, but that in certain areas, it is slightly inclined, or even opposite to the flow, in particular downstream of the dilution holes. . There may be detachments from the cooling air in these areas.
  • the purpose of the present invention is to prevent the air from the multi-perforation from taking off from the wall.
  • the present invention therefore proposes to locally orient the orifices as a function of the local flow of the burnt gases.
  • the combustion chamber is characterized in that the wall is subdivided into several zones, in each of which the orifices are defined by inclinations A and angles B having respectively identical values and calculated according to the characteristics of the flow of burnt gases in each of said zones.
  • Said wall is subdivided in particular into first zones situated respectively downstream of the dilution holes, and in which the orifices are directed against the current of the general direction of the flow of the burnt gases, of the second and third zones arranged on both sides another of said first zones with respect to the axial planes passing through the corresponding dilution holes, and a fourth zone covering the rest of said wall.
  • the orifices formed in the fourth zone have an axial inclination greater than 30. Their angle B is substantially equal to 0 °. The flow of fresh air from these orifices licks the internal surface of the wall in the direction of the axial flow of the burnt gases.
  • the orifices formed in the first zones diffuse a cooling air counter-current to the general direction of flow of the burnt gases.
  • Their inclination A is between 0 ° and -60 °, and their angle B is substantially equal to 0 °.
  • each of the first zones in the circumferential direction, there is provided a second and a third zone, the orifices of which distribute cooling air towards the axial plane passing through the corresponding dilution hole and in the direction of the general flow of burnt gases.
  • the combustion chamber 1 of annular type, comprises an outer annular axial wall 2 and an inner annular axial wall 3, joined at their upstream ends by a chamber bottom 4 equipped with injection systems 5, and having between their downstream ends an annular opening 6 for the exhaust of the burnt gases G towards a turbine not shown in the drawings.
  • the burnt gases G circulate in the internal cavity 7 of the combustion chamber 1 in a general axial direction represented by the arrow D.
  • outer 2 and inner 3 axial walls define, with the outer 8 and inner 9 casings, annular passages 10 and 11 in which circulates cooling air A coming from a compressor not shown in the drawings and located upstream of the combustion chamber 1.
  • the two walls 2 and 3 are provided with a plurality of dilution holes 12 regularly distributed in an axial plane 13 perpendicular to the axis of the turbomachine, and with a plurality of through holes 14 constituting a multi-perforation.
  • Part of the cooling air A enters axially into the internal cavity 7 through the dilution holes 12 and participates in the depletion and cooling of the combustion gases in the dilution zone of the combustion chamber 1, while the rest of the air A enters the internal cavity 7 through the orifices 14 in order to form a cooling film on the internal faces 2a and 3a of the axial walls 2 and 3.
  • FIG. 2 shows the diagram of the gas velocities in the vicinity of the internal face 2a of the external wall 2, in the region of two dilution holes 12a and 12b, this diagram having been obtained by 3D calculations.
  • This diagram shows that, in the zone 15 which separates the two dilution holes 12a and 12b, the gases flow in the direction D.
  • the gases on the contrary flow towards the dilution holes 12a and 12b, that is to say in a direction generally opposite to the direction D.
  • the gases flow in a direction inclined towards the axial plane 18 passing through the corresponding dilution hole, and generally directed in the general flow direction of the burnt gases D.
  • the burnt gases circulate in the direction D.
  • the 3D temperature diagram in the vicinity of the dilution holes also shows significant differences depending on the zones.
  • the region of the wall 2 and 3 which has the orifices 14 is subdivided into several zones, in each of which, the angles of inclination A of the axes 30 of the orifices 14 relative to the normal to the wall are identical , as well as the angles B of the planes 32 containing said axes 30 and the normals 31 with respect to the axial planes 33 containing said normals.
  • FIG 3 there is shown an axial wall portion 34 having two dilution holes 12a and 12b.
  • the arrow D represents the general direction of the flow of the burnt gases in the combustion chamber 1.
  • the references 16a and 16b represent the first zones in which the burnt gases flow against the current.
  • the burnt gases flow generally in the direction of the arrows 19.
  • the gases flow in the direction of the arrows 20.
  • the orifices 14 formed in the fourth zone 21 are defined by an inclination A 4 greater than 30 ° and an angle B substantially equal to 0 °.
  • the cooling air diffused by these orifices 14 enters the combustion chamber 1, in the general flow direction D of the gases, but with an inclination A 4 .
  • the orifices 14 formed in the first zone 16a are inclined so as to allow a diffusion of cooling air against the current of the general direction D.
  • the axes 30 of these orifices 14 form an angle A 1 with the normals 31 which is between -60 ° and 0 °.
  • the axes 30 of these orifices 14 are also parallel to the axial plane 18a passing through the axis 35 of the dilution hole 12a.
  • FIG 5 there is shown a small part 36 of the outer wall 2 at a third area 19b.
  • the orifices are drilled at an inclination A 3 relative to the normal 31 and in a plane making an angle B 3 with respect to the direction of the main flow D.
  • the angle B 3 is calculated in function of the average direction of local gas flow in the third zone 19b.

Landscapes

  • Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
EP96400863A 1995-04-26 1996-04-24 Brennkammer mit einer Vielzahl von Filmkühlungsbohrungen die in verschiedene axiale und tangentiale Richtungen geneigt sind Expired - Lifetime EP0743490B1 (de)

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
FR9504968A FR2733582B1 (fr) 1995-04-26 1995-04-26 Chambre de combustion comportant une multiperforation d'inclinaison axiale et tangentielle variable
FR9504968 1995-04-26

Publications (2)

Publication Number Publication Date
EP0743490A1 true EP0743490A1 (de) 1996-11-20
EP0743490B1 EP0743490B1 (de) 1999-06-09

Family

ID=9478445

Family Applications (1)

Application Number Title Priority Date Filing Date
EP96400863A Expired - Lifetime EP0743490B1 (de) 1995-04-26 1996-04-24 Brennkammer mit einer Vielzahl von Filmkühlungsbohrungen die in verschiedene axiale und tangentiale Richtungen geneigt sind

Country Status (5)

Country Link
US (1) US5775108A (de)
EP (1) EP0743490B1 (de)
JP (1) JP3302559B2 (de)
DE (1) DE69602804T2 (de)
FR (1) FR2733582B1 (de)

Cited By (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP1840466A1 (de) * 2006-03-30 2007-10-03 Snecma Anordnung von Verdünnungsöffnungen in der Brennkammerwand einer Turbomaschine
FR2974162A1 (fr) * 2011-04-14 2012-10-19 Snecma Virole de tube a flamme dans une chambre de combustion de turbomachine

Families Citing this family (39)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
FR2770283B1 (fr) * 1997-10-29 1999-11-19 Snecma Chambre de combustion pour turbomachine
US6145319A (en) * 1998-07-16 2000-11-14 General Electric Company Transitional multihole combustion liner
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
FR2856468B1 (fr) * 2003-06-17 2007-11-23 Snecma Moteurs Chambre de combustion annulaire de turbomachine
FR2856467B1 (fr) * 2003-06-18 2005-09-02 Snecma Moteurs Chambre de combustion annulaire de turbomachine
US7146816B2 (en) * 2004-08-16 2006-12-12 Honeywell International, Inc. Effusion momentum control
US20060037323A1 (en) * 2004-08-20 2006-02-23 Honeywell International Inc., Film effectiveness enhancement using tangential effusion
US7464554B2 (en) * 2004-09-09 2008-12-16 United Technologies Corporation Gas turbine combustor heat shield panel or exhaust panel including a cooling device
US7614235B2 (en) * 2005-03-01 2009-11-10 United Technologies Corporation Combustor cooling hole pattern
FR2892180B1 (fr) * 2005-10-18 2008-02-01 Snecma Sa Amelioration des perfomances d'une chambre de combustion par multiperforation des parois
US7631502B2 (en) * 2005-12-14 2009-12-15 United Technologies Corporation Local cooling hole pattern
US7546737B2 (en) * 2006-01-24 2009-06-16 Honeywell International Inc. Segmented effusion cooled gas turbine engine combustor
US7887322B2 (en) * 2006-09-12 2011-02-15 General Electric Company Mixing hole arrangement and method for improving homogeneity of an air and fuel mixture in a combustor
US7942006B2 (en) * 2007-03-26 2011-05-17 Honeywell International Inc. Combustors and combustion systems for gas turbine engines
US8091367B2 (en) * 2008-09-26 2012-01-10 Pratt & Whitney Canada Corp. Combustor with improved cooling holes arrangement
FR2941287B1 (fr) * 2009-01-19 2011-03-25 Snecma Paroi de chambre de combustion de turbomachine a une seule rangee annulaire d'orifices d'entree d'air primaire et de dilution
US8640464B2 (en) * 2009-02-23 2014-02-04 Williams International Co., L.L.C. Combustion system
FR2955374B1 (fr) * 2010-01-15 2012-05-18 Turbomeca Chambre de combustion multi-percee a ecoulements tangentiels contre giratoires
FR2979416B1 (fr) * 2011-08-26 2013-09-20 Turbomeca Paroi de chambre de combustion
US20160201908A1 (en) * 2013-08-30 2016-07-14 United Technologies Corporation Vena contracta swirling dilution passages for gas turbine engine combustor
US9453424B2 (en) * 2013-10-21 2016-09-27 Siemens Energy, Inc. Reverse bulk flow effusion cooling
FR3014115B1 (fr) 2013-12-02 2017-04-28 Office National Detudes Et De Rech Aerospatiales Onera Procede et systeme de depot d'oxyde sur un composant poreux
FR3013996B1 (fr) 2013-12-02 2017-04-28 Office National Detudes Et De Rech Aerospatiales Onera Procede de reparation locale de barrieres thermiques
WO2015103357A1 (en) 2013-12-31 2015-07-09 United Technologies Corporation Gas turbine engine wall assembly with enhanced flow architecture
EP3099976B1 (de) * 2014-01-30 2019-03-13 United Technologies Corporation Kühlfluss für führungspaneel in einer gasturbinenbrennkammer
US20160258623A1 (en) * 2015-03-05 2016-09-08 United Technologies Corporation Combustor and heat shield configurations for a gas turbine engine
DE102016201452A1 (de) 2016-02-01 2017-08-03 Rolls-Royce Deutschland Ltd & Co Kg Gasturbinenbrennkammer mit Wandkonturierung
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
US10823410B2 (en) 2016-10-26 2020-11-03 Raytheon Technologies Corporation Cast combustor liner panel radius for gas turbine engine combustor
US10669939B2 (en) 2016-10-26 2020-06-02 Raytheon Technologies Corporation Combustor seal for a gas turbine engine combustor
US10670269B2 (en) * 2016-10-26 2020-06-02 Raytheon Technologies Corporation Cast combustor liner panel gating feature for a gas turbine engine combustor
US10830448B2 (en) 2016-10-26 2020-11-10 Raytheon Technologies Corporation Combustor liner panel with a multiple of heat transfer augmentors for a gas turbine engine combustor
US10935243B2 (en) 2016-11-30 2021-03-02 Raytheon Technologies Corporation Regulated combustor liner panel for a gas turbine engine combustor
US11015529B2 (en) 2016-12-23 2021-05-25 General Electric Company Feature based cooling using in wall contoured cooling passage
US10480327B2 (en) 2017-01-03 2019-11-19 General Electric Company Components having channels for impingement cooling
US10753283B2 (en) 2017-03-20 2020-08-25 Pratt & Whitney Canada Corp. Combustor heat shield cooling hole arrangement
US11029027B2 (en) 2018-10-03 2021-06-08 Raytheon Technologies Corporation Dilution/effusion hole pattern for thick combustor panels
CN113251441B (zh) * 2021-06-28 2022-03-25 南京航空航天大学 一种新型航天发动机用多斜孔板椭球摆冷却结构

Citations (6)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
FR2410138A2 (fr) * 1977-11-29 1979-06-22 Snecma Perfectionnements aux chambres de combustion pour moteur a turbine a gaz
FR2635577A1 (fr) * 1988-08-17 1990-02-23 Rolls Royce Plc Chambre de combustion pour moteur a turbine a gaz
EP0512670A1 (de) * 1991-05-03 1992-11-11 General Electric Company Vorzugskühlungsmuster von Filmkühlungsbohrungen für Brennkammerwand
EP0592161A1 (de) * 1992-10-06 1994-04-13 ROLLS-ROYCE plc Gasturbinenbrennkammer
US5307637A (en) * 1992-07-09 1994-05-03 General Electric Company Angled multi-hole film cooled single wall combustor dome plate
US5323602A (en) * 1993-05-06 1994-06-28 Williams International Corporation Fuel/air distribution and effusion cooling system for a turbine engine combustor burner

Family Cites Families (12)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
FR512723A (fr) * 1920-01-05 1921-01-29 Rene Lucien Joseph Pierrel Générateur de gaz sous pression par combustion continue en vase clos
GB1271084A (en) * 1968-05-13 1972-04-19 T C Borrie Ltd Improvements in or relating to cartridge-operated hand tools
IL42390A0 (en) * 1972-08-02 1973-07-30 Gen Electric Impingement cooled combustor dome
US3916619A (en) * 1972-10-30 1975-11-04 Hitachi Ltd Burning method for gas turbine combustor and a construction thereof
DE2607214A1 (de) * 1976-02-23 1977-09-01 Volkswagenwerk Ag Brennkammer fuer gasturbinen
CH634128A5 (de) * 1978-06-13 1983-01-14 Bbc Brown Boveri & Cie Kuehlvorrichtung an einer wand.
US4790140A (en) * 1985-01-18 1988-12-13 Ishikawajima-Harima Jukogyo Kabushiki Kaisha Liner cooling construction for gas turbine combustor or the like
DE3711751A1 (de) * 1987-04-07 1988-10-20 Bergwerksverband Gmbh Gleichstromdatenuebertragungseinrichtung
US5329773A (en) * 1989-08-31 1994-07-19 Alliedsignal Inc. Turbine combustor cooling system
US5181379A (en) * 1990-11-15 1993-01-26 General Electric Company Gas turbine engine multi-hole film cooled combustor liner and method of manufacture
CA2056592A1 (en) * 1990-12-21 1992-06-22 Phillip D. Napoli Multi-hole film cooled combustor liner with slotted film starter
FR2714152B1 (fr) * 1993-12-22 1996-01-19 Snecma Dispositif de fixation d'une tuile de protection thermique dans une chambre de combustion.

Patent Citations (6)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
FR2410138A2 (fr) * 1977-11-29 1979-06-22 Snecma Perfectionnements aux chambres de combustion pour moteur a turbine a gaz
FR2635577A1 (fr) * 1988-08-17 1990-02-23 Rolls Royce Plc Chambre de combustion pour moteur a turbine a gaz
EP0512670A1 (de) * 1991-05-03 1992-11-11 General Electric Company Vorzugskühlungsmuster von Filmkühlungsbohrungen für Brennkammerwand
US5307637A (en) * 1992-07-09 1994-05-03 General Electric Company Angled multi-hole film cooled single wall combustor dome plate
EP0592161A1 (de) * 1992-10-06 1994-04-13 ROLLS-ROYCE plc Gasturbinenbrennkammer
US5323602A (en) * 1993-05-06 1994-06-28 Williams International Corporation Fuel/air distribution and effusion cooling system for a turbine engine combustor burner

Non-Patent Citations (2)

* Cited by examiner, † Cited by third party
Title
KÖHLER + BEER: "calculation of the disturbance to combustion chamber film cooling due to air injection through a row of jets", ZEITSCHRIFT FUR FLUGWISSENSCHAFTEN UND WELTRAUMFORSCHUNG, vol. 9, no. 1, February 1985 (1985-02-01), KOLN DE, pages 34 - 42, XP002010370 *
STEVENS S J ET AL: "EXPERIMENTAL STUDIES OF COMBUSTOR DILUTION ZONE AERODYNAMICS PART I: MEAN FLOWFIELDS", JOURNAL OF PROPULSION AND POWER, vol. 6, no. 3, 1 May 1990 (1990-05-01), pages 297 - 304, XP000126906 *

Cited By (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP1840466A1 (de) * 2006-03-30 2007-10-03 Snecma Anordnung von Verdünnungsöffnungen in der Brennkammerwand einer Turbomaschine
FR2899315A1 (fr) * 2006-03-30 2007-10-05 Snecma Sa Configuration d'ouvertures de dilution dans une paroi de chambre de combustion de turbomachine
US7891194B2 (en) 2006-03-30 2011-02-22 Snecma Configuration of dilution openings in a turbomachine combustion chamber wall
FR2974162A1 (fr) * 2011-04-14 2012-10-19 Snecma Virole de tube a flamme dans une chambre de combustion de turbomachine

Also Published As

Publication number Publication date
JP3302559B2 (ja) 2002-07-15
EP0743490B1 (de) 1999-06-09
US5775108A (en) 1998-07-07
DE69602804D1 (de) 1999-07-15
JPH08312960A (ja) 1996-11-26
DE69602804T2 (de) 2000-01-27
FR2733582B1 (fr) 1997-06-06
FR2733582A1 (fr) 1996-10-31

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