EP2031304B1 - Séparateur pour alimentation de l'air de refroidissement d'une turbine - Google Patents

Séparateur pour alimentation de l'air de refroidissement d'une turbine Download PDF

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Publication number
EP2031304B1
EP2031304B1 EP08163040.2A EP08163040A EP2031304B1 EP 2031304 B1 EP2031304 B1 EP 2031304B1 EP 08163040 A EP08163040 A EP 08163040A EP 2031304 B1 EP2031304 B1 EP 2031304B1
Authority
EP
European Patent Office
Prior art keywords
combustion chamber
separator
tubular portion
upstream
chamber
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
EP08163040.2A
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German (de)
English (en)
French (fr)
Other versions
EP2031304A1 (fr
Inventor
Christophe Pieussergues
Denis Sandelis
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 EP2031304A1 publication Critical patent/EP2031304A1/fr
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Publication of EP2031304B1 publication Critical patent/EP2031304B1/fr
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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
    • F23R3/42—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
    • 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 field of annular combustion chambers.
  • upstream and downstream are defined with respect to the direction of normal circulation of the air along the outside of the annular wall of the combustion chamber.
  • inner and outer / “outer” characterize a position more or less remote from the main axis of the combustion chamber, unless otherwise specified.
  • This combustion chamber is typically delimited by a bottom wall 12 comprising the fuel injectors 13 and the combustion air inlets, and by an annular wall 15 extending in the longitudinal direction of the chamber 10 (which therefore corresponds to the upstream-downstream direction), substantially parallel to the main axis A of the turbomachine (not shown).
  • the chamber 10 is closed at its upstream end by the bottom wall 12, and is open at its downstream end 17, in its longitudinal direction, to allow the evacuation of the burnt gases.
  • This annular wall 15 is typically constituted by an annular inner ring (radially inner wall) 151 and an annular outer ring (radially outer wall) 152.
  • the inner ferrule 151 and the outer ferrule 152 are coaxial with respect to the main axis A of the turbomachine, the inner ferrule 151 being closer to the main axis of the turbomachine than the outer ferrule 152, that is to say say having a radius less than the radius of the outer shell 152.
  • an upstream annular inner wall 11 of chamber 10 extends upstream the inner ring 151.
  • the annular wall 15 is pierced over its entire surface area (or a major part of it) with several orifices, larger or smaller, which are intended to allow air to enter the combustion chamber 10.
  • the air which runs along the inner ferrule 151 outside the chamber 10, and which then enters this chamber through these orifices flows between the inner ferrule 151 and a wall called internal chamber flange 21.
  • the internal flange 21 is pierced with orifices, some of which (upstream orifices 215) are located on its upstream part, substantially facing the central part of the inner ferrule 21 of the chamber 10 (that is, say halfway between the bottom wall 12 of the chamber 10 and the downstream end 217 of the inner flange 21).
  • upstream orifices 215 are located on its upstream part, substantially facing the central part of the inner ferrule 21 of the chamber 10 (that is, say halfway between the bottom wall 12 of the chamber 10 and the downstream end 217 of the inner flange 21).
  • the air flow intended to pass through the orifices of the inner flange to cool the HP turbine wheel is influenced by the combustion chamber. Indeed, this air is, before passing through these orifices, in contact with the inner wall which is hot and which is further pierced with air inlet orifices, and this air is thus subjected to heating by convection. This air also undergoes radiation heating through these openings of the chamber, this radiation from the flames of combustion. In addition, the instabilities of this combustion generate in this flow of air, through the orifices of the chamber, turbulence may contribute to disrupt the supply of cooling air to the wheel HP.
  • annular combustion chamber having the features according to the preamble of claim 1 is known from the document US 2002/108 374 .
  • the invention aims to provide a device that reduces the heating of the air for cooling the HP turbine wheel, and to reduce the disturbance of this air caused by the combustion instabilities from the combustion chamber.
  • the combustion chamber is equipped with a separator disposed between the radially inner wall of the chamber and the internal flange of the chamber, this separator comprising a tubular portion centered on the main axis of the chamber. of combustion and whose upstream end is located upstream of the orifices of the radially inner wall of the chamber, and a fastening portion integral with the combustion chamber, so that the tubular portion divides the air flow along this radially inner wall in an interior air flow passing between this tubular portion and the internal flange of the chamber, and in an outside air flow passing between the radially inner wall and this tubular portion.
  • the internal air flow which is intended to cool the HP turbine wheel, is no longer heated by convection and radiation by the chamber wall and by the flame radiation, and is no longer disturbed. by the instabilities of combustion resulting from the orifices of the internal wall of the cooling chamber.
  • the undesirable interaction between the combustion chamber and the air flow intended to cool the HP turbine wheel is greatly reduced or even eliminated.
  • the fixing portion is a radial portion extending from the tubular portion to the main axis, and is pierced with main holes for passing air from upstream to downstream.
  • the separator is not attached directly to the wall (hot) of the chamber, and is not heated by it by solid conduction. This arrangement is advantageous since the separator must be as warm as possible so as not to heat the interior air flow.
  • the figure 1 represents a combustion chamber 10 of a turbomachine and the structures associated therewith.
  • This room excluding the elements according to the invention, is identical to the chamber according to the art previous ( figure 5 ) described above. Common areas figures 1 and 5 therefore have the same numbering, and are not described again.
  • the downstream end of the outer shell 152 is extended radially outwardly by an annular outer flange 22, and the downstream end of the inner shell 151 is extended radially inwardly by an annular inner flange 21.
  • These flanges are therefore integral with the chamber 10.
  • the outer flange 22 and the inner flange 21 are attached to a housing wall 30 which surrounds the chamber 10, and thus serve to fix this chamber on the housing which is integral with the turbomachine.
  • the inner flange 21 extends the downstream end of the inner ferrule 151 inwards and then upstream, so that the inner flange 21, which is coaxial with the inner ferrule 151, has a radius smaller than that of this inner ferrule 151.
  • the inner flange 21 thus delimits with the inner ferrule 151 a downstream annular vein 40.
  • the upstream end 211 of the inner flange 21 is radial and is fixed (for example by several bolts / nuts distributed circumferentially along this upstream end 211), on a downstream end 301 radial wall of the casing 30.
  • the wall of casing 30 extends the inner flange 21 upstream, thus delimiting with the upstream annular inner wall 11 of the chamber 10 an upstream annular vein 49 (which extends downstream by the downstream annular vein 40).
  • the upstream end 211 of the inner flange 21 is in the longitudinal direction substantially at the upstream portion of the inner shell 151 (which terminates upstream approximately at the bottom wall 12 chamber). In the example shown in the figures, this upstream end 211 is located substantially at the first quarter upstream of the length between the bottom wall 12 and the downstream end 217 of the internal flange 21 (this downstream end 217 being located at the downstream end 17 of the chamber 10).
  • downstream annular groove 40 tapers from upstream to downstream, so that the radial dimension of the downstream annular groove 40 at the upstream end 211 of the inner flange 21 is greater than the radial dimension of the annular vein 40 at the downstream end 217 of the internal flange 21.
  • the inner flange 21 is pierced with orifices, including upstream orifices 215.
  • the portion of the air coming from the upstream annular duct 49 which passes through these upstream orifices 215 of the internal flange 21 is intended to go cool the HP turbine wheel (not shown). On the figure 1 this air passes, after having passed through the upstream orifices 215, through a structure 60 before going to cool this turbine.
  • a separator 70 is placed in the downstream annular groove 40, that is to say between the inner shroud 151 and the assembly formed by the internal flange 21 and the housing wall 30.
  • this separator 70 comprises a tubular portion 76 centered on the main axis A of the combustion chamber 10, and a radial portion 71 extending radially from the tubular portion 76 towards the main axis A, and pierced with main holes 72 oriented along the main axis A.
  • the radial portion 71 of the separator 70 is connected to the tubular portion 76 of the separator 70 in the upstream half of this tubular portion 76.
  • the radial portion 71 is connected to the tubular portion 76 at the first upstream quarter or the first third upstream of the tubular portion 76.
  • the tubular portion 76 of the separator 70 divides, from its upstream end 79, the downstream annular groove 40 in two in the upstream-downstream direction, firstly into an outer annular vein 81 located between the inner ferrule 151 of the chamber 10 and this tubular portion 76, and secondly an inner annular vein 82 located between this tubular portion 76 and the assembly constituted by the inner flange 21 and the housing wall 30. More specifically, the portion 78 of the tubular portion 76 located upstream of the radial portion 71 of the separator 70 is located between the housing wall 30 and the inner shell 151. The portion of the tubular portion 76 located downstream of the radial portion 71 is between the inner flange 21 and the inner ferrule 151.
  • the radial portion 71 of the separator 70 is thus located at the interface between the housing wall 30 and the internal flange 21.
  • the separator 70 is fixed to the internal flange 21 by the radially inner end of its radial portion 71.
  • the radially inner end of the radial portion 71 is pierced with fixing holes 711 adapted to receive a device for attaching said radial portion (71) to said inner flange (21).
  • this fixing is done by bolting.
  • the radially inner end of the radial portion 71 is sandwiched between the upstream radial end 211 of the inner flange 21 and the downstream radial end 301 of the casing wall 30.
  • the bolts that hold this upstream end 211 and this downstream end 301 integral pass through the fixing holes 711, the assembly consisting of the upstream end 211, the inner end of the radial portion 71, and the downstream end 301 being tightened by screwed nuts on these bolts.
  • the separator 70 is thus firmly held in position in the downstream annular groove 40.
  • the tubular portion 76 of the separator 70 divides the downstream annular vein 40 in the upstream-downstream direction into an outer annular vein 81 located between the inner ferrule 151 of the chamber 10 and this tubular portion 76, and in one Inner annular groove 82.
  • the tubular portion 76 does not have any holes, since its function is to separate the air circulating in the outer annular vein 81 (and which is heated by the chamber 10) of the air circulating in the annular vein Inner 82.
  • the tubular portion 76 screens between the air flowing in the inner annular vein 82 and the chamber 10.
  • the air coming from the upstream annular vein 49 thus divides into the downstream annular groove 40, at the upstream end 79 of the tubular portion 76 of the separator 70, into an outside air flow F e passing through the vein outer annulus 81, and an internal air flow F i passing into the inner annular vein 82 (these flows are represented by arrows on the figure 2 ).
  • the cross section (radial) of the outer annular vein 81 is smaller than the cross section of the downstream annular groove 40 in the absence of separator 70. Furthermore, the tubular portion 76 of the separator 70, and in particular its portion 78 located upstream of the radial portion 71 of the separator, is substantially parallel to the inner ring 151 of the combustion chamber 10.
  • the outer annular vein 81 is therefore of substantially constant cross section, which was not the case in the absence of separator 70, the inner flange 21 approaching the inner ferrule 151 from upstream to downstream.
  • This characteristic of the outer annular vein 81 leads to a better air flow, and thus to an increase in the Mach number in the outer annular vein 81.
  • This increase in the Mach number allows better cooling by convection of the inner ring 151 of the chamber 10.
  • the tests carried out by the inventors show that the increase in the Mach number is of the order of 10 to 20%.
  • the interior air flow F i flows between the casing wall 30 and the inner casing 151. Then it passes through the main holes 72 of the radial part 71 of the separator 70 and opens into the part of the inner vein 82 delimited by the inner flange 21 and the inner ferrule 151.
  • the separator 70 is in contact with the internal flange 21, so that the downstream end of the Inner annular vein 82 is closed.
  • the downstream end 77 is in contact with a portion 27 of the inner flange 21 which is an annular protrusion as shown in FIG. figure 2 .
  • the internal flange 21 has in its upstream portion upstream orifices 215. These upstream orifices 215 are located between the portion 27 and the upstream end 211 of the inner flange 21.
  • the interior air flow F i must therefore go through the upstream ports 215 of the inner flange 21 to exit the inner annular vein 82. This indoor air flow F i then flows in the direction of the HP turbine wheel that it is intended to go cool.
  • the downstream end 77 of the separator 70 can simply be slid over the portion 27 of the inner flange 21, which helps centering the separator 70 on the internal flange 21.
  • the downstream end 77 of the separator 70 may be attached to the portion 27 of the inner flange 21, for example by brazing.
  • the fixing is not done by bolting, which allows an easier assembly of the separator 70 on the inner flange 21.
  • the separator 70 is thus fixed to the inner flange 21 at the same time by the radially inner end of its part. radial 71, and the downstream end 77 of its tubular portion 76. This double attachment of the separator 70 allows a better attachment thereof on the inner flange 21.
  • the separator 70 is fixed to the inner flange 21 by its upstream and downstream, which improves the stability of the positioning of the separator 70, and stiffens the structure.
  • the separator 70 may comprise, instead of the radial portion 71, a fixing portion which is integral with the combustion chamber 10.
  • the separator 70 can be fixed rigidly (for example by welding) by the downstream end 77 of its tubular portion 76 on the inner flange 21 (for example on the portion 27 of the inner flange 21).
  • the separator 70 comprises only the tubular portion 76 and does not have a radial portion 71, and the downstream end 77 is the fixing portion.
  • the fixing portion can be connected to the tubular portion 76 in the upstream half of this tubular portion 76.
  • the upstream end 79 of the tubular portion 76 of the separator 70 is located upstream of the orifices of the inner ferrule 151 of the chamber 10. This situation is represented on the figure 2 , where the upstream end 79 is at a distance d upstream of the orifice 51 of the inner ferrule 151 situated furthest upstream. This distance d is for example between 15 and 20 mm.
  • the internal air flow F i is completely separated from the inner ferrule 151 of the chamber 10 by the tubular portion 76 of the separator 70.
  • the interior air flow F i is therefore not heated by convection at contact of the inner ring 151, or by the flame radiation passing through the orifices of the inner ring 151, and is also not disturbed by the instabilities of combustion from these orifices.
  • the indoor air flow F i can therefore cool more efficiently the HP turbine.
  • leading edge of the upstream end 79 of the tubular portion 76 of the separator 70 may be rounded, which allows a better flow of the outside air flow F e intended to follow the chamber 10 and the flow of the indoor air F i for cooling the HP turbine wheel.
  • the radial portion 71 of the separator has main holes 72 for passing the interior air flow F i .
  • These main holes 72 are located near the tubular portion 76, between this tubular portion 76 and the location where the radial portion 71 joins the inner flange 21.
  • main holes 72 are for example distributed over the entire circumference of the radial portion 71. They are for example circular, as shown in FIG. Figure 4A , or triangular arranged in staggered rows (i.e. any two adjacent triangles form a rhombus), as shown in FIG. Figure 4B .
  • the cross section of the radial portion 71 is defined as the region of this radial portion which is subjected to the interior air flow F i .
  • This effective section is therefore the annular region between the place where the radial portion 71 is connected to the tubular portion 76 (this location is substantially a circle in the example shown in the figures), and the location where the radial portion 71 comes into contact with the inner flange 21 (this location is substantially a circle in the example shown in the figures).
  • the area of the main holes 72 occupies between 60% and 80% of the effective cross section of the radial portion 71.
  • the separator material is capable of holding temperatures up to 550 ° C.
  • this material may be a nickel / chromium steel.
  • the combustion chamber described above is a turbomachine chamber. This chamber can also be any combustion chamber.

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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)
EP08163040.2A 2007-08-31 2008-08-27 Séparateur pour alimentation de l'air de refroidissement d'une turbine Active EP2031304B1 (fr)

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
FR0757283A FR2920525B1 (fr) 2007-08-31 2007-08-31 Separateur pour alimentation de l'air de refroidissement d'une turbine

Publications (2)

Publication Number Publication Date
EP2031304A1 EP2031304A1 (fr) 2009-03-04
EP2031304B1 true EP2031304B1 (fr) 2015-11-18

Family

ID=39327261

Family Applications (1)

Application Number Title Priority Date Filing Date
EP08163040.2A Active EP2031304B1 (fr) 2007-08-31 2008-08-27 Séparateur pour alimentation de l'air de refroidissement d'une turbine

Country Status (6)

Country Link
US (1) US8069669B2 (ru)
EP (1) EP2031304B1 (ru)
JP (1) JP5384052B2 (ru)
CA (1) CA2639178C (ru)
FR (1) FR2920525B1 (ru)
RU (1) RU2477822C2 (ru)

Families Citing this family (5)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP1892378A1 (de) * 2006-08-22 2008-02-27 Siemens Aktiengesellschaft Gasturbine
FR2953907B1 (fr) * 2009-12-11 2012-11-02 Snecma Chambre de combustion pour turbomachine
US20150241067A1 (en) * 2012-09-26 2015-08-27 United Technologies Corporation Fastened joint for a tangential on board injector
US9476429B2 (en) * 2012-12-19 2016-10-25 United Technologies Corporation Flow feed diffuser
US11371700B2 (en) * 2020-07-15 2022-06-28 Raytheon Technologies Corporation Deflector for conduit inlet within a combustor section plenum

Citations (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US20020108374A1 (en) * 2001-02-09 2002-08-15 Young Craig Douglas Slot cooled combustor liner

Family Cites Families (35)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2938342A (en) * 1954-08-24 1960-05-31 Rolls Royce Gas turbine engines
US3228190A (en) * 1962-08-16 1966-01-11 Power Jets Res & Dev Ltd Gas turbine plant
US3797236A (en) * 1971-06-11 1974-03-19 Rolls Royce Annular combustion chamber with ceramic annular ring
US3989410A (en) * 1974-11-27 1976-11-02 General Electric Company Labyrinth seal system
US3986720A (en) * 1975-04-14 1976-10-19 General Electric Company Turbine shroud structure
USH903H (en) * 1982-05-03 1991-04-02 General Electric Company Cool tip combustor
US4466239A (en) * 1983-02-22 1984-08-21 General Electric Company Gas turbine engine with improved air cooling circuit
US5211536A (en) * 1991-05-13 1993-05-18 General Electric Company Boltless turbine nozzle/stationary seal mounting
FR2691235B1 (fr) 1992-05-13 1995-07-07 Snecma Chambre de combustion comprenant un ensemble separateur des gaz.
DE59208715D1 (de) * 1992-11-09 1997-08-21 Asea Brown Boveri Gasturbinen-Brennkammer
FR2706021B1 (fr) 1993-06-03 1995-07-07 Snecma Chambre de combustion comprenant un ensemble séparateur de gaz.
FR2706534B1 (fr) 1993-06-10 1995-07-21 Snecma Diffuseur-séparateur multiflux avec redresseur intégré pour turboréacteur.
FR2712379B1 (fr) 1993-11-10 1995-12-29 Snecma Chambre de combustion pour turbomachine munie d'un séparateur des gaz.
FR2721694B1 (fr) 1994-06-22 1996-07-19 Snecma Refroidissement de l'injecteur de décollage d'une chambre de combustion à deux têtes.
FR2723177B1 (fr) 1994-07-27 1996-09-06 Snecma Chambre de combustion comportant une double paroi
RU2107227C1 (ru) * 1995-11-01 1998-03-20 Акционерное общество "Авиадвигатель" Трубчато-кольцевая камера сгорания газотурбинной энергетической установки
RU2109218C1 (ru) * 1996-02-06 1998-04-20 Акционерное общество "Авиадвигатель" Трубчато-кольцевая камера сгорания газовой турбины
FR2752916B1 (fr) 1996-09-05 1998-10-02 Snecma Chemise de protection thermique pour chambre de combustion de turboreacteur
FR2758384B1 (fr) 1997-01-16 1999-02-12 Snecma Controle des debits de refroidissement pour des chambres de combustion a haute temperature
US6266961B1 (en) * 1999-10-14 2001-07-31 General Electric Company Film cooled combustor liner and method of making the same
US6540477B2 (en) * 2001-05-21 2003-04-01 General Electric Company Turbine cooling circuit
FR2825787B1 (fr) * 2001-06-06 2004-08-27 Snecma Moteurs Montage de chambre de combustion cmc de turbomachine par viroles de liaison souples
US6553767B2 (en) * 2001-06-11 2003-04-29 General Electric Company Gas turbine combustor liner with asymmetric dilution holes machined from a single piece form
JP3924136B2 (ja) * 2001-06-27 2007-06-06 三菱重工業株式会社 ガスタービン燃焼器
FR2840974B1 (fr) 2002-06-13 2005-12-30 Snecma Propulsion Solide Anneau d'etancheite pour cahmbre de combustion et chambre de combustion comportant un tel anneau
JP3840556B2 (ja) * 2002-08-22 2006-11-01 川崎重工業株式会社 燃焼器ライナのシール構造
US7093441B2 (en) * 2003-10-09 2006-08-22 United Technologies Corporation Gas turbine annular combustor having a first converging volume and a second converging volume, converging less gradually than the first converging volume
FR2866079B1 (fr) 2004-02-05 2006-03-17 Snecma Moteurs Diffuseur pour turboreacteur
FR2880391A1 (fr) 2005-01-06 2006-07-07 Snecma Moteurs Sa Diffuseur pour chambre annulaire de combustion, en particulier pour un turbomoteur d'avion
FR2885201B1 (fr) 2005-04-28 2010-09-17 Snecma Moteurs Chambre de combustion aisement demontable a performance aerodynamique amelioree
FR2893390B1 (fr) 2005-11-15 2011-04-01 Snecma Fond de chambre de combustion avec ventilation
FR2897107B1 (fr) 2006-02-09 2013-01-18 Snecma Paroi transversale de chambre de combustion munie de trous de multiperforation
FR2897418B1 (fr) * 2006-02-10 2013-03-01 Snecma Chambre de combustion annulaire d'une turbomachine
FR2903172B1 (fr) 2006-06-29 2008-10-17 Snecma Sa Agencement pour chambre de combustion de turbomachine ayant un defecteur a collerette
FR2909748B1 (fr) 2006-12-07 2009-07-10 Snecma Sa Fond de chambre, procede de realisation de celui-ci, chambre de combustion le comportant et turboreacteur en etant equipe

Patent Citations (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US20020108374A1 (en) * 2001-02-09 2002-08-15 Young Craig Douglas Slot cooled combustor liner

Also Published As

Publication number Publication date
JP5384052B2 (ja) 2014-01-08
CA2639178A1 (fr) 2009-02-28
US8069669B2 (en) 2011-12-06
RU2008135300A (ru) 2010-03-10
JP2009057970A (ja) 2009-03-19
RU2477822C2 (ru) 2013-03-20
FR2920525B1 (fr) 2014-06-13
EP2031304A1 (fr) 2009-03-04
CA2639178C (fr) 2016-02-09
US20090060723A1 (en) 2009-03-05
FR2920525A1 (fr) 2009-03-06

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