EP2488792B1 - Mehrpunkteinspritzer für eine brennkammer eines turbinenmotors - Google Patents

Mehrpunkteinspritzer für eine brennkammer eines turbinenmotors Download PDF

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Publication number
EP2488792B1
EP2488792B1 EP10779566.8A EP10779566A EP2488792B1 EP 2488792 B1 EP2488792 B1 EP 2488792B1 EP 10779566 A EP10779566 A EP 10779566A EP 2488792 B1 EP2488792 B1 EP 2488792B1
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EP
European Patent Office
Prior art keywords
annular
fuel
chamber
ring
orifices
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Application number
EP10779566.8A
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English (en)
French (fr)
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EP2488792A1 (de
Inventor
Didier Hippolyte Hernandez
Thomas Olivier Marie Noel
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Safran Aircraft Engines SAS
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SNECMA SAS
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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
    • F23R3/28—Continuous combustion chambers using liquid or gaseous fuel characterised by the fuel supply
    • F23R3/283—Attaching or cooling of fuel injecting means including supports for fuel injectors, stems, or lances
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23D—BURNERS
    • F23D11/00—Burners using a direct spraying action of liquid droplets or vaporised liquid into the combustion space
    • F23D11/36—Details, e.g. burner cooling means, noise reduction means
    • 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
    • F23R3/28—Continuous combustion chambers using liquid or gaseous fuel characterised by the fuel supply
    • F23R3/34—Feeding into different combustion zones
    • F23R3/343—Pilot flames, i.e. fuel nozzles or injectors using only a very small proportion of the total fuel to insure continuous combustion
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23D—BURNERS
    • F23D2900/00—Special features of, or arrangements for burners using fluid fuels or solid fuels suspended in a carrier gas
    • F23D2900/00016—Preventing or reducing deposit build-up on burner parts, e.g. from carbon

Definitions

  • the present invention relates to a "multipoint" fuel injection device for an annular turbomachine combustion chamber such as an airplane turbojet or turboprop engine.
  • a turbomachine comprises an annular combustion chamber arranged at the outlet of a high pressure compressor and provided with a plurality of fuel injection devices regularly distributed circumferentially at the inlet of the combustion chamber.
  • Each multipoint injection device comprises a first venturi inside which is mounted a pilot injector centered on the axis of the first venturi and continuously supplied by a pilot circuit and a second venturi coaxial with the first venturi and surrounding it.
  • This second venturi comprises an annular chamber at its upstream end in which is mounted an annular ring supplied with fuel by a multipoint circuit.
  • the ring has fuel injection orifices formed in a front face facing downstream and outward of the second venturi.
  • the pilot circuit continuously provides optimized fuel flow for low revs and the multi-point circuit provides optimized intermittent fuel flow for high revs.
  • Such a fuel injection device is known from the document EP 1 806 536 of the Applicant.
  • the intermittent use of the multipoint circuit has the major drawback of inducing, under the effect of the high temperatures due to the radiation of the flame in the combustion chamber, a scrub or a coking of the stagnant fuel inside the circuit. multipoint when it is cut off. These phenomena can lead to formation of coke in the ring and at the fuel injection ports of the multipoint circuit impacting the fuel spraying by the multipoint circuit and therefore the operation of the combustion chamber.
  • Such a configuration does not, however, sufficiently reduce the risk of coking fuel circulating at the front face of the annular chamber which remains highly exposed to the heat radiation generated by the combustion of the fuel downstream.
  • the invention aims in particular to provide a simple, effective and economical solution to this problem.
  • a fuel injection device for an annular turbomachine combustion chamber comprising a pilot circuit continuously supplying an injector opening into a first venturi and a multipoint circuit supplying intermittently injection orifices formed in a front face of an upstream annular chamber of a second venturi coaxial with and surrounding the first venturi, an annular ring being mounted in the annular chamber to delimit a fuel supply circuit of the injection ports and a cooling circuit by passage of fuel supplying the injector of the pilot circuit, characterized in that the cooling circuit extends on the front face of the chamber in the immediate vicinity of the injection orifices.
  • a part of the cooling circuit is formed by a groove of a downstream face of the annular ring, this downstream face being applied to the front face of the annular chamber.
  • the cooling circuit also comprises an annular channel formed between the inner cylindrical walls of the ring and the annular chamber, in order to cool the internal cylindrical face of the annular chamber of the second venturi inside which circulates a flow of hot air from the high pressure compressor.
  • the cooling circuit further comprises an annular channel formed between the outer cylindrical walls of the annular ring and the annular chamber, this channel being able to be used for cooling the outer wall of the annular chamber by circulating the fuel of the pilot circuit or for being intended to be isolated from the pilot circuit and to be filled in operation with air or coked fuel as thermal insulation.
  • the outer periphery of the annular chamber of the second venturi is subjected to temperatures lower than those of the inner periphery of the annular chamber and it is therefore not necessary to continuously cool the outer contour of the annular chamber and the use of thermal insulation proves sufficient.
  • the cooling circuit of the front face of the chamber is corrugated and extends alternately radially inside and outside of the injection orifices, which makes it possible to position the cooling system as close as possible to the injection ports.
  • the cooling circuit of the front face of the chamber comprises two semicircular symmetrical branches each extending between input means and fuel outlet means, the latter being connected to the injector of the pilot circuit.
  • the injection of fuel through the orifices of the annular chamber is performed through orifices of the crown which open into the orifices of the annular chamber.
  • the orifices of the downstream wall of the ring have a smaller diameter than that of the orifices of the end face of the annular chamber, which prevents fuel drops coming out of the orifices of the ring from coking the holes of the ring. the wall of the chamber, when stopping the multipoint circuit.
  • the invention also relates to an annular turbomachine combustion chamber comprising at least one fuel injection device of the type described above.
  • the invention also relates to a turbomachine, such as a turbojet engine or a turboprop engine, comprising at least one fuel injection device of the type described above.
  • FIG. 1 representing an injection device 10 comprising two fuel injection systems, one of which is a permanently operating pilot system and the other a multipoint system operating intermittently.
  • This device is intended to be mounted in an opening of a bottom wall of an annular combustion chamber of a turbomachine which is supplied with air by an upstream high-pressure compressor and whose combustion gases feed a turbine mounted downstream.
  • This device comprises a first venturi 12 and a second venturi 14 coaxial, the first venturi 12 being mounted inside the second venturi 14.
  • a pilot injector 16 is mounted inside a first stage of tendrils 18 inserted axially to Inside the first venturi 12.
  • a second stage of tendrils 20 is formed at the upstream end and radially outside the first venturi 12 and separates the first 12 and second 14 venturis.
  • the second venturi 14 comprises an annular chamber 22 formed by two radially inner cylindrical walls 24 and outer 26 connected to each other by a frustoconical downstream wall 28 converging downstream.
  • An annular ring 30 also comprising two radially inner cylindrical walls 32 and external wall 34 connected to each other by a downstream convergent downstream conical downstream wall 36 is mounted inside the annular chamber 22 so that the downstream walls 28, 36 of the annular chamber 22 and the annular ring 30 are in contact.
  • the annular ring 30 is centered inside the annular chamber 22 by means of an annular shoulder 38 formed inside the annular chamber 30 at the junction of the frustoconical downstream wall 28 and the internal cylindrical wall 24 of the chamber ring 22.
  • the annular ring 30 and the annular chamber 22 each comprise an annular opening at their upstream end.
  • the cylindrical walls 24, 26 of the annular chamber 22 extend projecting upstream with respect to the upstream ends of the cylindrical walls 32, 34 of the annular ring 30.
  • the downstream wall 36 of the annular ring 30 comprises injection orifices 40 uniformly distributed circumferentially and opening into corresponding orifices 42 in the downstream wall 28 of the chamber 22.
  • the orifices 40, 42 of the annular chamber 22 and the annular ring 30 have identical diameters.
  • An inner annular channel 44 is defined between the inner cylindrical walls 24, 32 of the annular ring 30 and the annular chamber 22.
  • an outer annular channel 46 is defined between the outer cylindrical walls 26, 34 of the annular ring. 30 and the annular chamber 22.
  • the injection device comprises a body 48 whose downstream portion is annular and comprises a cylindrical duct 50 axially engaged sealing between the inner cylindrical walls 24 and outer 26 of the annular chamber 22 and opening sealing between the inner cylindrical walls 32 and outer ring 34 of the annular ring 30.
  • the duct 50 has a radial shoulder 54 abutting on the upstream ends of the inner cylindrical walls 32 and outer 34 of the annular ring 30.
  • This sealing assembly of the body 48 makes it possible to guarantee that the inner and outer annular channels 44 and 44 are sealed with respect to the annular space formed inside the annular ring 30.
  • a fuel supply arm 56 is connected to the body 48 and comprises two coaxial ducts of which one central 58 feeds a channel 60 of the body 48 opening downstream inside the annular ring 30 and the other formed outer 62 around the central duct 58 supplies at the output of separate channels (not shown) opening into the inner annular channels 44 and outer 46, respectively.
  • the body 48 comprises a fuel collection cavity 64 formed diametrically opposite the fuel supply arm 56 and at the upstream ends of the cylindrical walls 32, 34 of the annular ring 30 so that the annular channels internal 44 and outer 46 communicate with the collection cavity 64.
  • a duct 66 is connected at one end to the pilot injector 16 and at the other end to the body 48 and opens into the collection cavity 64.
  • the central duct 58 of the arm 56 supplies fuel to the channel 60 of the body 48, the fuel then circulating in the annular ring 30 and being injected into the combustion chamber downstream through the orifices 40, 42 of the ring gear 30 and from bedroom 22.
  • the outer conduit 62 of the arm 56 feeds the channels of the body 48 opening into the inner annular channels 44 and outer 46, the fuel then flowing into the collection cavity 64 to supply the pilot injector 16 via the conduit 66 .
  • This circuit forms the pilot circuit and operates continuously while the multipoint circuit operates intermittently during specific flight phases such as takeoff requiring additional power.
  • the hot air (at about 600 ° C.) coming from the high-pressure compressor flows inside the first venturi 12, in the first radial swirler 18, and from the air also flows inside the second radial swirler 20, between the first 12 and second 14 venturis.
  • the front downstream face 28 of the annular chamber 22 is also subjected to thermal radiation combustion, which can lead to coking of the fuel in the injection orifices 40, 42 of the ring 30 and the chamber annular 22 during flight phases where the multipoint circuit is not used.
  • the invention provides a solution to this problem by integrating in the injection device 67 a cooling circuit of the frustoconical front wall 68 of the annular chamber 70, in the immediate vicinity of the injection orifices as shown in the drawings. Figures 2 to 4 .
  • This cooling circuit comprises a groove 72 formed on the downstream face of the frustoconical wall 74 of the annular ring 76, which is applied to the upstream face of the frustoconical wall 68 of the annular chamber 70.
  • the groove 72 is corrugated and extends alternately radially inside and inside the injection ports 78 of the annular ring 76, which makes it possible to cool as much as possible the orifices 78 of the ring 76 and the orifices 80 of the annular chamber 70.
  • the groove 72 comprises two semicircular branches fueled by two channels 82, 84 of the body 48 and connected to the output cavity 64 diametrically opposite. The two branches are symmetrical with respect to a plane passing through the axis of the pilot injector 16 and halfway between the two channels 82, 84 for feeding the groove 72.
  • the cooling circuit according to the invention also comprises an internal annular groove 86 formed in the thickness of the inner cylindrical wall 88 of the ring 76, this groove 86 delimiting an internal annular channel with the internal cylindrical wall 90 of the annular chamber 70
  • the inner annular channel is supplied with fuel by two channels 92, 94 of the body 48 and is connected at the outlet to the collection cavity 64, for cooling the internal cylindrical walls 88, 90 of the annular ring 76 and the annular chamber 70.
  • Two semicircular grooves 96, 98 are formed in the thickness of the outer cylindrical wall 100 of the annular ring 76 and delimit with the outer cylindrical wall 102 of the annular chamber 70 two semicircular channels whose circumferential ends are closed by axial ribs 104 of the annular ring 76. In this way, the two outer semi-circular channels are isolated from the collection chamber supplying the pilot injector.
  • the two channels 96, 98 semicircular are filled with air.
  • these channels can be filled with air if the seal is realized with respect to the pilot circuit and in particular with respect to the front circuit or they can be filled with fuel in the opposite case, which fuel cokes under the effect of the high temperatures.
  • the air or the coked fuel forms a thermal insulator, which is sufficient to prevent coking of the fuel inside the ring as the outer peripheries of the annular ring 76 and the annular chamber 70 are subjected to temperatures lower than those to which the internal peripheries of these same parts are subjected.
  • the orifices 78 of the downstream frustoconical wall 74 of the annular ring 76 have a diameter smaller than that of the orifices of the frustoconical front face 68 of the annular chamber 70. This prevents, when the multipoint circuit is stopped, that drops of fuel remain at the orifices 78 of the annular ring 76 do not coke the orifices 80 of the annular chamber 70.
  • the diameter of the orifices 78 of the annular ring 76 is of the order of 0.5 mm and that of the orifices 80 of the annular chamber 70 is of the order of 1 mm.
  • the downstream face of the frustoconical wall 74 of the ring 72 is fixed to seal on the frustoconical wall 68 of the annular chamber 70, for example by brazing.
  • the junction between an orifice 78 of the ring 76 and an orifice 80 of the annular chamber 70 is sealed.
  • brazing it is possible to produce in one piece, for example by laser sintering, the annular ring 76 and the second venturi 14 comprising the annular chamber 70.
  • the invention is not limited to the corrugated cooling circuit as described above. It is thus possible to form two grooves in the downstream face of the downstream wall 74 of the ring 76, one of the grooves being located radially inside the orifices 78 of the ring 76 and the other being located radially at the outside of these same orifices 78.
  • this circuit would not better cool the orifices 78, 80 of the annular ring 76 and the annular chamber 70 and in particular the circumferential inter-orifices spaces.
  • the external channels 96, 98 are connected to the collection cavity 64 supplying the pilot injector 16 and participate in the cooling of the annular chamber 70 by circulating the fuel of the pilot injector 16.

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  • Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Fuel-Injection Apparatus (AREA)

Claims (12)

  1. Vorrichtung zum Einspritzen von Kraftstoff für eine ringförmige Verbrennungskammer einer Turbomaschine bzw. eines Turbotriebwerks, enthaltend einen Pilotkreis, der einen Injektor (16) dauerhaft speist, welcher in eine erste Venturi-Düse (12) mündet, und einen Mehrpunktkreis, der Einspritzöffnungen (80) intermittierend speist, die in einer Stirnseite (68) einer stromaufwärtigen, ringförmigen Kammer (70) einer zweiten Venturi-Düse (14) ausgeführt sind, welche koaxial zur ersten Venturi-Düse (12) verläuft und diese umgibt, wobei ein ringförmiger Kranz (76) in der ringförmigen Kammer (70) gelagert ist, um darin einen Kraftstoffversorgungskreis für die Einspritzöffnungen (80) und einen Durchlaufkühlkreis für den Kraftstoff zu begrenzen, mit dem der Injektor des Pilotkreises gespeist wird, dadurch gekennzeichnet, dass der Kühlkreis sich über die Stirnseite (68) der Kammer (70) in unmittelbarer Nähe zu den Einspritzöffnungen (80) erstreckt.
  2. Vorrichtung nach Anspruch 1, dadurch gekennzeichnet, dass der Kühlkreis eine in einer stromabwärtigen Seite des ringförmigen Kranzes (76) ausgebildete Nut aufweist, wobei diese stromabwärtige Seite an der Stirnseite (68) der ringförmigen Kammer (70) anliegt.
  3. Vorrichtung nach einem der Ansprüche 1 oder 2, dadurch gekennzeichnet, dass der Kühlkreis auch einen ringförmigen Kanal aufweist, der zwischen den zylindrischen Innenwänden (88, 90) des Kranzes (76) und der ringförmigen Kammer (70) ausgebildet ist.
  4. Vorrichtung nach einem der Ansprüche 1 bis 3, dadurch gekennzeichnet, dass der Kühlkreis auch einen ringförmigen Kanal aufweist, der zwischen den zylindrischen Außenwänden (100, 102) des Kranzes (76) und der ringförmigen Kammer (70) ausgebildet ist.
  5. Vorrichtung nach Anspruch 4, dadurch gekennzeichnet, dass der zwischen den zylindrischen Außenwänden (100, 102) des Kranzes (76) und der ringförmigen Kammer (70) ausgebildete ringförmige Kanal dazu bestimmt ist, vom Pilotkreis getrennt zu werden und im Betrieb mit Luft oder verkoktem Kraftstoff gefüllt zu werden.
  6. Vorrichtung nach einem der Ansprüche 1 bis 5, dadurch gekennzeichnet, dass der Kühlkreis der Stirnseite (68) der Kammer (70) gewellt ausgeführt ist und sich abwechselnd radial innerhalb und außerhalb der Einspritzöffnungen (80) erstreckt.
  7. Vorrichtung nach einem der Ansprüche 1 bis 6, dadurch gekennzeichnet, dass der Kühlkreis der Stirnseite (68) der Kammer (70) zwei symmetrische, halbrunde Zweige aufweist, die sich jeweils zwischen Kraftstoffeinlassmitteln und Kraftstoffauslassmitteln erstrecken.
  8. Vorrichtung nach Anspruch 7, dadurch gekennzeichnet, dass die Kraftstoffauslassmittel mit dem Injektor (16) des Pilotkreises verbunden sind.
  9. Vorrichtung nach einem der Ansprüche 1 bis 8, dadurch gekennzeichnet, dass die stromabwärtige Wand (74) des Kranzes (76) Durchtrittsöffnungen (78) für den Durchtritt von Kraftstoff aufweist, die in die vorgenannten Öffnungen (80) der Stirnseite (68) der ringförmigen Kammer (70) münden.
  10. Vorrichtung nach Anspruch 9, dadurch gekennzeichnet, dass die Öffnungen (78) der stromabwärtigen Wand (74) des Kranzes (76) einen Durchmesser haben, der kleiner als der der Öffnungen (80) der Stirnwand (68) der ringförmigen Kammer (70) ist.
  11. Ringförmige Verbrennungskammer einer Turbomaschine bzw. eines Turbotriebwerks, dadurch gekennzeichnet, dass sie zumindest eine Vorrichtung (67) zum Einspritzen von Kraftstoff nach einem der vorangehenden Ansprüche aufweist.
  12. Turbomaschine bzw. Turbinentriebwerk, wie etwa Turbostrahltriebwerk oder Turbopropellertriebwerk, dadurch gekennzeichnet, dass sie bzw. es eine Vorrichtung (67) zum Einspritzen von Kraftstoff nach einem der Ansprüche 1 bis 10 aufweist.
EP10779566.8A 2009-10-13 2010-10-12 Mehrpunkteinspritzer für eine brennkammer eines turbinenmotors Active EP2488792B1 (de)

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
FR0904907A FR2951246B1 (fr) 2009-10-13 2009-10-13 Injecteur multi-point pour une chambre de combustion de turbomachine
PCT/FR2010/000682 WO2011045486A1 (fr) 2009-10-13 2010-10-12 Injecteur multi-point pour une chambre de combustion de turbomachine

Publications (2)

Publication Number Publication Date
EP2488792A1 EP2488792A1 (de) 2012-08-22
EP2488792B1 true EP2488792B1 (de) 2015-03-25

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EP10779566.8A Active EP2488792B1 (de) 2009-10-13 2010-10-12 Mehrpunkteinspritzer für eine brennkammer eines turbinenmotors

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US (1) US9046271B2 (de)
EP (1) EP2488792B1 (de)
JP (1) JP5762424B2 (de)
CN (1) CN102575844B (de)
BR (1) BR112012008441B1 (de)
CA (1) CA2776843C (de)
FR (1) FR2951246B1 (de)
RU (1) RU2543097C2 (de)
WO (1) WO2011045486A1 (de)

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US9267689B2 (en) * 2013-03-04 2016-02-23 Siemens Aktiengesellschaft Combustor apparatus in a gas turbine engine
FR3003632B1 (fr) * 2013-03-19 2016-10-14 Snecma Systeme d'injection pour chambre de combustion de turbomachine comportant une paroi annulaire a profil interne convergent
EP3033508B1 (de) 2013-08-16 2018-06-20 United Technologies Corporation Gekühltes kraftstoffeinspritzsystem für einen gasturbinenmotor
US9556795B2 (en) * 2013-09-06 2017-01-31 Delavan Inc Integrated heat shield
FR3011318B1 (fr) * 2013-10-01 2018-01-05 Safran Aircraft Engines Injecteur de carburant dans une turbomachine
US10012197B2 (en) 2013-10-18 2018-07-03 Holley Performance Products, Inc. Fuel injection throttle body
FR3017416B1 (fr) * 2014-02-12 2018-12-07 Safran Aircraft Engines Refroidissement d'une canalisation principale dans un systeme carburant a injection multipoints
CN105650678B (zh) * 2016-01-11 2018-04-10 清华大学 涡轮活塞混合动力系统的燃烧室进气结构
US9376997B1 (en) 2016-01-13 2016-06-28 Fuel Injection Technology Inc. EFI throttle body with side fuel injectors
US20240271571A1 (en) * 2023-02-14 2024-08-15 Collins Engine Nozzles, Inc. Proportional control of cooling circuit of fuel nozzle
US20240271790A1 (en) * 2023-02-14 2024-08-15 Collins Engine Nozzles, Inc. Variable cooling of secondary circuit of fuel nozzles

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FR2596102B1 (fr) * 1986-03-20 1988-05-27 Snecma Dispositif d'injection a vrille axialo-centripete
FR2673705A1 (fr) * 1991-03-06 1992-09-11 Snecma Chambre de combustion de turbomachine munie d'un dispositif anti-cokefaction du fond de ladite chambre.
US6389815B1 (en) * 2000-09-08 2002-05-21 General Electric Company Fuel nozzle assembly for reduced exhaust emissions
FR2832493B1 (fr) * 2001-11-21 2004-07-09 Snecma Moteurs Systeme d'injection multi-etages d'un melange air/carburant dans une chambre de combustion de turbomachine
US6898938B2 (en) * 2003-04-24 2005-05-31 General Electric Company Differential pressure induced purging fuel injector with asymmetric cyclone
FR2896031B1 (fr) * 2006-01-09 2008-04-18 Snecma Sa Dispositif d'injection multimode pour chambre de combustion, notamment d'un turboreacteur
FR2896030B1 (fr) * 2006-01-09 2008-04-18 Snecma Sa Refroidissement d'un dispositif d'injection multimode pour chambre de combustion, notamment d'un turboreacteur
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FR2919898B1 (fr) * 2007-08-10 2014-08-22 Snecma Injecteur multipoint pour turbomachine
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Publication number Publication date
CN102575844B (zh) 2014-12-31
BR112012008441A2 (pt) 2016-03-29
US9046271B2 (en) 2015-06-02
FR2951246B1 (fr) 2011-11-11
CA2776843A1 (fr) 2011-04-21
BR112012008441B1 (pt) 2020-09-29
US20120198852A1 (en) 2012-08-09
RU2543097C2 (ru) 2015-02-27
RU2012119573A (ru) 2013-11-20
CA2776843C (fr) 2017-07-04
CN102575844A (zh) 2012-07-11
EP2488792A1 (de) 2012-08-22
WO2011045486A1 (fr) 2011-04-21
FR2951246A1 (fr) 2011-04-15
JP5762424B2 (ja) 2015-08-12
JP2013507599A (ja) 2013-03-04

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