US20150354505A1 - Propellant feed circuit and a cooling method - Google Patents

Propellant feed circuit and a cooling method Download PDF

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Publication number
US20150354505A1
US20150354505A1 US14/760,175 US201414760175A US2015354505A1 US 20150354505 A1 US20150354505 A1 US 20150354505A1 US 201414760175 A US201414760175 A US 201414760175A US 2015354505 A1 US2015354505 A1 US 2015354505A1
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US
United States
Prior art keywords
feed circuit
propellant
tank
heat exchanger
heat
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.)
Abandoned
Application number
US14/760,175
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English (en)
Inventor
Didier Vuillamy
Gérard ROZ
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.)
ArianeGroup SAS
Original Assignee
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 SNECMA SAS filed Critical SNECMA SAS
Assigned to SNECMA reassignment SNECMA ASSIGNMENT OF ASSIGNORS INTEREST (SEE DOCUMENT FOR DETAILS). Assignors: ROZ, Gérard, VUILLAMY, DIDIER
Publication of US20150354505A1 publication Critical patent/US20150354505A1/en
Assigned to AIRBUS SAFRAN LAUNCHERS SAS reassignment AIRBUS SAFRAN LAUNCHERS SAS ASSIGNMENT OF ASSIGNORS INTEREST (SEE DOCUMENT FOR DETAILS). Assignors: SNECMA
Assigned to ARIANEGROUP SAS reassignment ARIANEGROUP SAS CHANGE OF NAME (SEE DOCUMENT FOR DETAILS). Assignors: AIRBUS SAFRAN LAUNCHERS SAS
Abandoned legal-status Critical Current

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Classifications

    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02K—JET-PROPULSION PLANTS
    • F02K9/00—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof
    • F02K9/42—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof using liquid or gaseous propellants
    • F02K9/60—Constructional parts; Details not otherwise provided for
    • F02K9/605—Reservoirs
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02K—JET-PROPULSION PLANTS
    • F02K9/00—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof
    • F02K9/08—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof using solid propellants
    • F02K9/32—Constructional parts; Details not otherwise provided for
    • F02K9/40—Cooling arrangements
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02K—JET-PROPULSION PLANTS
    • F02K9/00—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof
    • F02K9/42—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof using liquid or gaseous propellants
    • F02K9/44—Feeding propellants
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02K—JET-PROPULSION PLANTS
    • F02K9/00—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof
    • F02K9/42—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof using liquid or gaseous propellants
    • F02K9/44—Feeding propellants
    • F02K9/46—Feeding propellants using pumps
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02K—JET-PROPULSION PLANTS
    • F02K9/00—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof
    • F02K9/42—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof using liquid or gaseous propellants
    • F02K9/60—Constructional parts; Details not otherwise provided for
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02K—JET-PROPULSION PLANTS
    • F02K9/00—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof
    • F02K9/42—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof using liquid or gaseous propellants
    • F02K9/60—Constructional parts; Details not otherwise provided for
    • F02K9/62—Combustion or thrust chambers
    • F02K9/64—Combustion or thrust chambers having cooling arrangements
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02K—JET-PROPULSION PLANTS
    • F02K9/00—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof
    • F02K9/97—Rocket nozzles
    • F02K9/972—Fluid cooling arrangements for nozzles
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2260/00—Function
    • F05D2260/20—Heat transfer, e.g. cooling
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2260/00—Function
    • F05D2260/20—Heat transfer, e.g. cooling
    • F05D2260/205—Cooling fluid recirculation, i.e. after cooling one or more components is the cooling fluid recovered and used elsewhere for other purposes

Definitions

  • the present invention relates to the aerospace field, and in particular to the field of vehicles propelled by rocket engines.
  • upstream and downstream are defined relative to the normal flow direction of propellants in the feed circuits of a rocket engine.
  • the present invention seeks to remedy those drawbacks.
  • the invention seeks to propose a feed circuit for feeding a rocket engine with at least a first liquid propellant, which circuit also serves to cool at least one heat source.
  • the feed circuit includes at least one buffer tank for said first liquid propellant and a first heat exchanger incorporated in said buffer tank and connected to a cooling circuit for cooling the at least one heat source.
  • the heat generated by the heat source can be removed via the cooling circuit and said first heat exchanger to the liquid propellant in the feed circuit of the rocket engine.
  • this cooling thus takes place via the intermediary of a cooling circuit interposed between the heat source and the propellant flowing through the feed circuit, thereby potentially enabling the temperature of the heat source to be regulated more accurately by the potential for regulating the flow rate of cooling fluid in the cooling circuit.
  • Incorporating the first heat exchanger in a buffer tank of the feed circuit makes it possible to increase the heat power that is absorbed, even when the feed circuit is off and said first propellant is not flowing.
  • the present description also relates to the assembly comprising said feed circuit and the heat source provided with a cooling circuit connected to said first heat exchanger of the feed circuit.
  • the heat source may in particular be a fuel cell.
  • a fuel cell may be fed with the same propellants as the rocket engine in order to generate electricity for on-board systems of a vehicle propelled by the rocket engine.
  • other types of on-board heat source e.g. such as batteries or electronic circuits, could nevertheless be cooled in the same manner.
  • the present invention also relates to a vehicle comprising a rocket engine with said feed circuit and an on-board heat generating device with a cooling circuit connected to said first heat exchanger of the feed circuit.
  • This vehicle may, for example, be a stage of a space launcher, a satellite, or any other type of vehicle that is propelled by a liquid propellant rocket engine.
  • said first liquid propellant may in particular be a cryogenic liquid, and in particular liquid hydrogen, thus providing cooling that is even more effective because of its low temperature.
  • said feed circuit may include a pump upstream from said first heat exchanger, in order to make the first propellant flow.
  • This pump may, for example, be an electric pump or a turbopump. Nevertheless, the feed circuit could alternatively be configured in such a manner as to make the first propellant flow by other means, e.g. such as by pressurizing a tank upstream.
  • the heated first propellant downstream from the first heat exchanger may be used not only for feeding the thrust chamber of the rocket engine or, possibly, a gas generator or the heat source itself (e.g. when the heat source is a fuel cell), but it may also be used in the gaseous state for maintaining the internal pressure in at least one tank of the first propellant while the tank is emptying through the feed circuit.
  • the feed circuit may include a branch leading to a high portion of this tank for the first propellant. The propellant in the gaseous state can thus be reinjected into the tank in order to maintain the internal pressure therein while the tank is emptying.
  • said feed circuit may include a branch passing through a second heat exchanger.
  • the second heat exchanger can thus allow a flow of the first propellant diverted through said branch to pass into the gaseous state, even when the heat power of said heat generating device, on its own, is insufficient for that purpose.
  • This flow of gas can thus be used, by way of example, for maintaining the internal pressure of a tank supplying the feed circuit with the first propellant as it empties.
  • the present description also relates to the assembly comprising the feed circuit and a tank for said first liquid propellant, the tank being connected to the feed circuit upstream from said first heat exchanger, and also to said branch downstream from said second heat exchanger.
  • said second heat exchanger may be incorporated in a tank for a second liquid propellant so as to be capable of heating the first liquid propellant by transferring heat from the second liquid propellant.
  • the second liquid propellant presents a boiling point that is significantly higher than the first liquid propellant (for example when the first liquid propellant is liquid hydrogen and the second liquid propellant is liquid oxygen)
  • this makes it possible not only to ensure that the first propellant passes into the gaseous phase in the second heat exchanger, but also, simultaneously, to cool the second propellant.
  • This cooling of the second propellant makes it possible to avoid cavitation in a pump downstream from the second tank.
  • the present description thus also relates to an assembly of this feed circuit and a tank for a second liquid propellant, and containing said second heat exchanger.
  • the present description also relates to a method of cooling a heat source, in which a cooling circuit of said heat source transfers the heat generated by the heat source to a first liquid propellant of a rocket engine via a first heat exchanger of a feed circuit for feeding said rocket engine at least with said first liquid propellant.
  • this first heat exchanger is contained in a buffer tank of the feed circuit for feeding the first propellant, and the heat source may be a fuel cell.
  • a portion of the flow of the first liquid propellant can then be diverted through a second heat exchanger in which it absorbs heat from a second propellant so as to reach the gaseous state prior to being injected into a tank for the first propellant feeding the feed circuit.
  • FIG. 1 is a diagrammatic view of a vehicle in a first embodiment of the invention
  • FIG. 2 is a diagrammatic view of a vehicle in a second embodiment of the invention.
  • FIG. 3 is a diagrammatic view of a vehicle in a third embodiment of the invention.
  • FIG. 1 is a diagram showing a vehicle 1 , which may for example be a stage of a space launcher.
  • this vehicle 1 has a liquid propellant rocket-engine 2 with a first tank 3 for a first propellant, a second tank 4 for a second propellant, a thrust chamber 5 for combustion of a mixture of the two propellants and for accelerating the gas that results from combustion of the mixture, a first feed circuit 6 connected to the first tank 3 and to the first chamber 5 in order to bring the first propellant from the first tank 3 to the thrust chamber 5 , and a second feed circuit 7 connected to the second tank 4 and to the thrust chamber 5 in order to bring the second propellant from the second tank 4 to the thrust chamber 5 .
  • the first and second propellants may be cryogenic propellants such as liquid hydrogen and liquid oxygen.
  • Each of the feed circuits 6 , 7 comprises a pump 8 , 9 for causing the respective propellant to flow through each feed circuit 6 , 7 , and outlet valves 10 , 11 in order to open and close the flow of propellants to the thrust chamber 5 .
  • these pumps 8 , 9 may be electric pumps, or they may be turbopumps.
  • the vehicle 1 also has an on-board fuel cell 16 adapted to generate electricity as a result of a chemical reaction between the two propellants, which fuel cell is connected to feed circuits 12 , 13 in order to be fed with these two propellants.
  • Each of these circuits 12 , 13 includes a micro-pump 14 , 15 for controlling the flow rate of fuel supplied to the fuel cell 16 .
  • the micro-pumps 14 , 15 could possibly be replaced by variable flow rate valves, with the internal pressure of the tanks 3 , 4 then sufficing to cause the propellants to flow towards the fuel cell 16 .
  • the fuel cell 16 is also provided with a cooling circuit 17 containing a cooling fluid such as, for example, helium and connected to a heat exchanger 18 incorporated in a buffer tank 20 of the feed circuit 6 for the first propellant.
  • a cooling circuit 17 containing a cooling fluid such as, for example, helium and connected to a heat exchanger 18 incorporated in a buffer tank 20 of the feed circuit 6 for the first propellant.
  • the flow of this cooling circuit in the cooling circuit 17 may be driven by, and may be regulated by means of a variable flow rate forced flow device 19 , which device is in the form of a fan in the embodiment shown.
  • the cooling fluid could be driven by a thermosiphon, and its flow rate could be regulated by at least one variable flow rate valve.
  • the pumps 8 , 9 drive the propellants via the feed circuits 6 , 7 to feed the thrust chamber 5 .
  • the heat generated by the fuel cell 16 which is fed simultaneously with propellants via the feed circuits 12 , 13 in order to generate electricity, is removed via the cooling circuit 17 and the heat exchanger 18 to the first propellant flowing through the feed circuit 6 .
  • the very low temperature of this first propellant when it is a cryogenic liquid, enables this heat to be removed extremely effectively.
  • a volume V t of 30 liters (L) of liquid hydrogen in the buffer tank 20 can thus absorb the quantity of heat that corresponds to thermal power P c of 100 watts (W) for one hour with a temperature rise AT of only 17 kelvins (K) in the liquid hydrogen.
  • a vehicle 1 in a second embodiment is shown in FIG. 2 .
  • This other vehicle 1 differs from the vehicle of the first embodiment in that the first feed circuit 6 includes, downstream from the buffer tank 20 , a return branch 21 returning to the top of the first tank 3 via a variable flow rate valve 22 , and a second heat exchanger 23 that is incorporated in the base of the second tank 4 in the proximity of its connection to the second feed circuit 7 .
  • the second circuit 7 Downstream from the pump 9 , the second circuit 7 also has a return branch 40 returning to the top of the second tank 4 , and passing through another heat exchanger 41 arranged around the thrust chamber 5 so as to be heated thereby by means of radiation or conduction.
  • this branch 40 also includes a valve 42 , which may be a variable flow rate valve, thereby enabling the flow rate through the branch 40 to be regulated accurately.
  • a valve 42 which may be a variable flow rate valve, thereby enabling the flow rate through the branch 40 to be regulated accurately.
  • a portion of the flow of the first propellant leaving the first tank 3 through the first feed circuit 6 is diverted through the branch 21 to the second heat exchanger 23 , in which it absorbs additional heat power from the higher-temperature second propellant, thereby passing into the gaseous state, prior to being injected into the top of the first tank 3 so as to maintain its internal pressure while it is emptying.
  • the temperature difference between their respective boiling points at atmospheric temperature is nearly 70 K, thus enabling a more than adequate quantity of heat to be transferred for vaporizing the liquid hydrogen before their temperatures become equal, with this applying even when the liquid hydrogen is flowing at a high rate relative to the volume of liquid oxygen contained in the second tank.
  • this absorption of heat by the second propellant in the second heat exchanger 23 cools the second propellant, thereby enabling the saturation pressure of the second propellant being fed to the pump 9 to be reduced so as to reduce cavitation phenomena in the pump. This also makes it possible to allow the pressure and the temperature of the second propellant to fluctuate more widely in the second tank 4 .
  • a portion of the flow of the second propellant extracted from the second tank 4 via the second circuit 7 is diverted through the branch 40 and is heated in the heat exchanger 41 the by heat radiation from the thrust chamber 5 , or by heat conduction, so that it passes into the gaseous phase prior to being reinjected into the second tank 4 , in order to maintain the internal pressure therein.
  • This diversion of flow is controlled by the valve 42 .
  • the flow of the propellants to the thrust chamber can also be provided by other means, for example such as pressurizing the tanks.
  • these pumps are replaced by a tank 24 of pressurized gas, e.g. helium, connected to the propellant tanks 3 and 4 via respective valves 26 and 27 .
  • pressurized gas e.g. helium
  • Pressurizing the propellants in the tanks 3 , 4 also makes it possible to omit micro-pumps for feeding the fuel cell 16 with propellants, with this feed being regulated in this embodiment by variable flow rate valves 28 , 29 in the circuits 12 , 13 .
  • the first feed circuit 6 includes a buffer tank 20 , and downstream therefrom it has a return branch 21 returning to the top of the first tank 3 via a variable flow rate valve 22 and a second heat exchanger 23 that is incorporated in the base of the second tank 4 in the proximity of its connection to the second feed circuit 7 , thereby making it possible to reduce the consumption of pressurized gas from the tank 24 for the purpose of pressurizing the first propellant tank 3 .
  • this branch 21 includes a forced flow device 30 , more specifically in the form of a fan or a pump.
  • the other elements of this vehicle 1 are essentially equivalent to elements of the second embodiment, and they are given the same reference numbers.
  • the vehicle could also have a branch for injecting the second propellants in the gaseous phase into the second tank, as in the second embodiment, including a device for forced flow of the second propellants in the gaseous phase. Consequently, the description and the drawings should be considered in a sense that is illustrative rather than restrictive.

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  • Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Fuel Cell (AREA)
  • Cooling, Air Intake And Gas Exhaust, And Fuel Tank Arrangements In Propulsion Units (AREA)
  • Electric Propulsion And Braking For Vehicles (AREA)
  • Filling Or Discharging Of Gas Storage Vessels (AREA)
US14/760,175 2013-01-11 2014-01-10 Propellant feed circuit and a cooling method Abandoned US20150354505A1 (en)

Applications Claiming Priority (3)

Application Number Priority Date Filing Date Title
FR1350239 2013-01-11
FR1350239A FR3000995B1 (fr) 2013-01-11 2013-01-11 Circuit d'alimentation en ergol et procede de refroidissement
PCT/FR2014/050045 WO2014108649A1 (fr) 2013-01-11 2014-01-10 Circuit d'alimentation en ergol et procede de refroidissement

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US14/760,190 Expired - Fee Related US10082106B2 (en) 2013-01-11 2014-01-10 Propellant feed circuit and a cooling method
US14/760,175 Abandoned US20150354505A1 (en) 2013-01-11 2014-01-10 Propellant feed circuit and a cooling method

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US (2) US10082106B2 (ru)
EP (2) EP2943675B1 (ru)
JP (2) JP6352306B2 (ru)
FR (3) FR3000995B1 (ru)
RU (2) RU2647353C2 (ru)
WO (3) WO2014108649A1 (ru)

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FR3000995B1 (fr) * 2013-01-11 2015-07-24 Snecma Circuit d'alimentation en ergol et procede de refroidissement
US10940961B2 (en) * 2015-01-14 2021-03-09 Ventions, Llc Small satellite propulsion system
CN104976929B (zh) * 2015-05-11 2017-05-31 上海宇航系统工程研究所 一种模拟气源装置
US10654592B2 (en) * 2016-02-12 2020-05-19 The Boeing Company Integration of fuel cell with cryogenic source for cooling and reactant
US10347923B2 (en) * 2016-03-04 2019-07-09 Teledyne Energy Systems, Inc. Fuel cell systems and cooling methods
FR3059092B1 (fr) * 2016-11-18 2018-12-14 Safran Aircraft Engines Dispositif pyrotechnique
FR3068082B1 (fr) * 2017-06-22 2019-08-09 Airbus Safran Launchers Sas Reservoir ameliore pour moteur d'engin spatial
CN109538378A (zh) * 2019-01-07 2019-03-29 西安交通大学 一种发动机燃气射流尾焰的处理系统
DE102019130787B4 (de) * 2019-11-14 2023-02-16 Deutsches Zentrum für Luft- und Raumfahrt e.V. Triebwerkanordnung, Verfahren zum Betreiben einer Triebwerkanordnung und Verwendung einer Durchflussbatterieanordnung bei einer Triebwerkanordnung
CN112431692B (zh) * 2020-11-17 2021-08-03 中国人民解放军战略支援部队航天工程大学 一种协同吸气式液体火箭发动机推进剂供应系统

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US3143855A (en) * 1959-06-30 1964-08-11 United Aircraft Corp Pressure fed propellant system for storable liquid rocket
US5644920A (en) * 1995-09-25 1997-07-08 Rockwell International Corporation Liquid propellant densification
US20020139902A1 (en) * 2001-03-16 2002-10-03 Snecma Moteurs Low-thrust cryogenic propulsion module
US6581882B2 (en) * 2001-03-16 2003-06-24 Snecma Moteurs Low-thrust cryogenic propulsion module
US20030005708A1 (en) * 2001-05-22 2003-01-09 Philip Beck Airborne gas storage and supply system
US6769242B1 (en) * 2001-11-21 2004-08-03 Mse Technology Applications, Inc. Rocket engine
US7418814B1 (en) * 2005-06-30 2008-09-02 The United States Of America As Represented By The Administrator Of The National Aeronautics And Space Administration Dual expander cycle rocket engine with an intermediate, closed-cycle heat exchanger
US7784269B1 (en) * 2006-08-25 2010-08-31 Xcor Aerospace System and method for cooling rocket engines
US20110005193A1 (en) * 2009-07-07 2011-01-13 Thomas Clayton Pavia Method and apparatus for simplified thrust chamber configurations
US9446862B2 (en) * 2011-06-17 2016-09-20 Snecma Cryogenic thruster assembly using regenerative heating from main and settling thrusters
US10082106B2 (en) * 2013-01-11 2018-09-25 Arianegroup Sas Propellant feed circuit and a cooling method

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WO2014108650A1 (fr) 2014-07-17
US10082106B2 (en) 2018-09-25
FR3000995A1 (fr) 2014-07-18
JP6352306B2 (ja) 2018-07-04
WO2014108651A2 (fr) 2014-07-17
RU2642711C2 (ru) 2018-01-25
FR3000995B1 (fr) 2015-07-24
RU2015133553A (ru) 2017-02-14
FR3000998A1 (fr) 2014-07-18
EP2943675B1 (fr) 2020-05-13
EP2943674A1 (fr) 2015-11-18
US20150337763A1 (en) 2015-11-26
RU2015133524A (ru) 2017-02-17
JP2016505764A (ja) 2016-02-25
FR3000997B1 (fr) 2015-06-05
EP2943675A1 (fr) 2015-11-18
FR3000997A1 (fr) 2014-07-18
RU2647353C2 (ru) 2018-03-15
JP6352305B2 (ja) 2018-07-04
WO2014108649A1 (fr) 2014-07-17
WO2014108651A3 (fr) 2014-11-13
JP2016505765A (ja) 2016-02-25

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