WO2014099085A2 - Turbine à gaz à bras de levier avant - Google Patents

Turbine à gaz à bras de levier avant Download PDF

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Publication number
WO2014099085A2
WO2014099085A2 PCT/US2013/062104 US2013062104W WO2014099085A2 WO 2014099085 A2 WO2014099085 A2 WO 2014099085A2 US 2013062104 W US2013062104 W US 2013062104W WO 2014099085 A2 WO2014099085 A2 WO 2014099085A2
Authority
WO
WIPO (PCT)
Prior art keywords
engine
fan blades
gas turbine
gravity
center
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.)
Ceased
Application number
PCT/US2013/062104
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English (en)
Other versions
WO2014099085A3 (fr
Inventor
David BOMZER
Frederick M. Schwarz
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.)
RTX Corp
Original Assignee
United Technologies Corp
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 United Technologies Corp filed Critical United Technologies Corp
Priority to US14/432,289 priority Critical patent/US11635025B2/en
Priority to CA2884976A priority patent/CA2884976C/fr
Priority to BR112015006820-0A priority patent/BR112015006820B1/pt
Priority to JP2015533317A priority patent/JP6027250B2/ja
Priority to EP13865250.8A priority patent/EP2904210A4/fr
Publication of WO2014099085A2 publication Critical patent/WO2014099085A2/fr
Publication of WO2014099085A3 publication Critical patent/WO2014099085A3/fr
Anticipated expiration legal-status Critical
Ceased legal-status Critical Current

Links

Classifications

    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D25/00—Component parts, details, or accessories, not provided for in, or of interest apart from, other groups
    • F01D25/28—Supporting or mounting arrangements, e.g. for turbine casing
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
    • F02C7/00—Features, components parts, details or accessories, not provided for in, or of interest apart form groups F02C1/00 - F02C6/00; Air intakes for jet-propulsion plants
    • F02C7/20—Mounting or supporting of plant; Accommodating heat expansion or creep
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02K—JET-PROPULSION PLANTS
    • F02K3/00—Plants including a gas turbine driving a compressor or a ducted fan
    • F02K3/02—Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber
    • F02K3/04—Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber the plant including ducted fans, i.e. fans with high volume, low pressure outputs, for augmenting the jet thrust, e.g. of double-flow type
    • F02K3/06—Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber the plant including ducted fans, i.e. fans with high volume, low pressure outputs, for augmenting the jet thrust, e.g. of double-flow type with front fan
    • 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
    • F05D2260/40—Transmission of power
    • F05D2260/403—Transmission of power through the shape of the drive components
    • F05D2260/4031—Transmission of power through the shape of the drive components as in toothed gearing
    • Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
    • Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
    • Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
    • Y02T50/00—Aeronautics or air transport
    • Y02T50/60—Efficient propulsion technologies, e.g. for aircraft

Definitions

  • a gas turbine engine typically includes a fan section, a compressor section, a combustor section and a turbine section. Air entering the compressor section is compressed and delivered into the combustion section where it is mixed with fuel and ignited to generate a high-speed exhaust gas flow. The high-speed exhaust gas flow expands through the turbine section to drive the compressor and the fan section.
  • the compressor section typically includes low and high pressure compressors, and the turbine section includes low and high pressure turbines.
  • a speed reduction device such as an epicyclical gear assembly may be utilized to drive the fan section such that the fan section may rotate at a speed different than the turbine section so as to increase the overall propulsive efficiency of the engine.
  • a shaft driven by one of the turbine sections provides an input to the epicyclical gear assembly that drives the fan section at a reduced speed such that both the turbine section and the fan section can rotate at closer to optimal speeds.
  • Structures of a gas turbine engine contribute to an overall weight of the engine and balance point is defined at center of gravity.
  • the location of the center of gravity of a gas turbine engine influences how an engine is mounted and how surrounding nacelle structures are configured.
  • a center of gravity moved forward along an engine axis increases an internal moment arm and thereby increases load on engine mounting structures.
  • the location of the engine center of gravity is influenced by selections of materials and component configurations.
  • a gas turbine engine includes a plurality of fan blades rotatable about an axis, each of the plurality of fan blades including a leading edge, a turbine section including an aft most turbine blade having a trailing edge, and a geared architecture driven by the turbine section for rotating the plurality of fan blades about the axis, wherein a center of gravity of the gas turbine engine is located a first axial distance from the trailing edge of the aft most turbine blade that is between about 35% and about 75% of a total length between the leading edge of the plurality of fan blades and the trailing edge of the aft most turbine blade.
  • the center of gravity is disposed substantially along the axis.
  • the center of gravity is determined including weights of structures comprising the gas turbine engine not including engine mounting structures, engine cowling structures and nacelle structures.
  • the center of gravity is determined including weights of fluids contained within operating systems of the gas turbine engine.
  • the first axial distance is between about 40% and about 70% of the total length between the leading edge of the plurality of fan blades and the trailing edge of the aft most turbine blade.
  • the plurality of fan blades are supported on a rotor with the rotor and fan blades having a density of between about 0.0094 lbs/ lbs/in 3 and about 0.01540 lbs/ lbs/in 3 .
  • the geared architecture comprises a gearbox having a density of between about 0.22 lbs/in 3 and about 0.30 lbs/in 3 .
  • the center gravity is located at an intersection of a vertical line extending through a hoist point of the gas turbine engine and the axis with the axis normal to the vertical line.
  • Another gas turbine engine includes a plurality of fan blades rotatable about an axis, each of the plurality of fan blades including a leading edge, a turbine section including an aft most rotating turbine blade having a trailing edge, and a geared architecture driven by the turbine section for rotating the plurality of fan blades about the axis, wherein an internal moment arm of the turbofan engine comprises a ratio of first distance from a center of gravity of the turbofan engine to the trailing edge of the aft most rotating turbine blade to a total length between the leading edge of the plurality of fan blades and the trailing edge of the aft most turbine blade that is between about 35% and about 75%.
  • the ratio is between about 40% and 70%.
  • the geared architecture comprises gearbox having a density of between about 0.22 lbs/in 3 and about 0.30 lbs/in 3 .
  • the center of gravity is determined including weights of fluids contained within operating systems of the turbofan engine.
  • the center gravity is located at an intersection of a vertical line extending through a hoist point of the turbofan engine and the axis with the axis normal to the vertical line.
  • a method of assembling a gas turbine engine includes supporting a plurality of fan blades about an axis of rotation with each of the plurality of fan blades including a leading edge, supporting a turbine section including an aft most turbine blade having a trailing edge about the axis of rotation, supporting a geared architecture driven by the turbine section for rotating the plurality of fan blades about the axis, and selecting components of the gas turbine engine structure to orientate a center of gravity of the gas turbine engine located a first axial distance from the trailing edge of the aft most turbine blade that is between about 35% and about 75% of a total length between the leading edge of the plurality of fan blades and the trailing edge of the aft most turbine blade.
  • a geared architecture as a gearbox having a density of between about 0.22 lbs/ lbs/in 3 and about 0.30 lbs/in 3 .
  • any of the foregoing methods including selecting components of the gas turbine engine to orientate the center of gravity within a range of between about 40% and 70% of total length between the leading edge of the plurality of fan blades and the trailing edge of the aft most turbine blade.
  • Figure 1 is a schematic view of a center of gravity of an example gas turbine engine.
  • Figure 2 is a schematic view of a center of gravity of an example gas turbine engine. DETAILED DESCRIPTION
  • FIG. 1 schematically illustrates an example gas turbine engine 20 that includes a fan section 22, a compressor section 24, a combustor section 26 and a turbine section 28.
  • Alternative engines might include an augmenter section (not shown) among other systems or features.
  • the fan section 22 drives air along a bypass flow path B while the compressor section 24 draws air in along a core flow path C where air is compressed and communicated to a combustor section 26.
  • air is mixed with fuel and ignited to generate a high pressure exhaust gas stream that expands through the turbine section 28 where energy is extracted and utilized to drive the fan section 22 and the compressor section 24.
  • turbofan gas turbine engine depicts a turbofan gas turbine engine
  • the concepts described herein are not limited to use with turbofans as the teachings may be applied to other types of turbine engines; for example a turbine engine including a three- spool architecture in which three spools concentrically rotate about a common axis and where a low spool enables a low pressure turbine to drive a fan via a gearbox, an intermediate spool that enables an intermediate pressure turbine to drive a first compressor of the compressor section, and a high spool that enables a high pressure turbine to drive a high pressure compressor of the compressor section.
  • the example engine 20 generally includes a low speed spool 30 and a high speed spool 32 mounted for rotation about an engine central longitudinal axis A relative to an engine static structure 36 via several bearing systems 38. It should be understood that various bearing systems 38 at various locations may alternatively or additionally be provided.
  • the low speed spool 30 generally includes an inner shaft 40 that connects a fan 42 and a low pressure (or first) compressor section 44 to a low pressure (or first) turbine section 46.
  • the inner shaft 40 drives the fan 42 through a speed change device, such as a geared architecture 48, to drive the fan 42 at a lower speed than the low speed spool 30.
  • the high-speed spool 32 includes an outer shaft 50 that interconnects a high pressure (or second) compressor section 52 and a high pressure (or second) turbine section 54.
  • the inner shaft 40 and the outer shaft 50 are concentric and rotate via the bearing systems 38 about the engine central longitudinal axis A.
  • a combustor 56 is arranged between the high pressure compressor 52 and the high pressure turbine 54.
  • the high pressure turbine 54 includes at least two stages to provide a double stage high pressure turbine 54. In another example, the high pressure turbine 54 includes only a single stage. As used herein, a "high pressure" compressor or turbine experiences a higher pressure than a corresponding "low pressure” compressor or turbine.
  • the example low pressure turbine 46 has a pressure ratio that is greater than about 5.
  • the pressure ratio of the example low pressure turbine 46 is measured prior to an inlet of the low pressure turbine 46 as related to the pressure measured at the outlet of the low pressure turbine 46 prior to an exhaust nozzle.
  • a mid-turbine frame 58 of the engine static structure 36 is arranged generally between the high pressure turbine 54 and the low pressure turbine 46.
  • the mid- turbine frame 58 further supports bearing systems 38 in the turbine section 28 as well as setting airflow entering the low pressure turbine 46.
  • Airflow through the core flow path C is compressed by the low pressure compressor 44 then by the high pressure compressor 52 mixed with fuel and ignited in the combustor 56 to produce high speed exhaust gases that are then expanded through the high pressure turbine 54 and low pressure turbine 46.
  • the mid-turbine frame 58 includes vanes 60, which are in the core airflow path and function as an inlet guide vane for the low pressure turbine 46. Utilizing the vane 60 of the mid-turbine frame 58 as the inlet guide vane for low pressure turbine 46 decreases the length of the low pressure turbine 46 without increasing the axial length of the mid-turbine frame 58. Reducing or eliminating the number of vanes in the low pressure turbine 46 shortens the axial length of the turbine section 28. Thus, the compactness of the gas turbine engine 20 is increased and a higher power density may be achieved.
  • the disclosed gas turbine engine 20 in one example is a high-bypass geared aircraft engine.
  • the gas turbine engine 20 includes a bypass ratio greater than about six (6), with an example embodiment being greater than about ten (10).
  • the example geared architecture 48 is an epicyclical gear train, such as a planetary gear system, star gear system or other known gear system, with a gear reduction ratio of greater than about 2.3.
  • the gas turbine engine 20 includes a bypass ratio greater than about ten (10: 1) and the fan diameter is significantly larger than an outer diameter of the low pressure compressor 44. It should be understood, however, that the above parameters are only exemplary of one embodiment of a gas turbine engine including a geared architecture and that the present disclosure is applicable to other gas turbine engines.
  • a significant amount of thrust is provided by the bypass flow B due to the high bypass ratio.
  • the fan section 22 of the engine 20 is designed for a particular flight condition - typically cruise at about 0.8 Mach and about 35,000 feet.
  • the flight condition of 0.8 Mach and 35,000 ft., with the engine at its best fuel consumption - also known as "bucket cruise Thrust Specific Fuel Consumption ('TSFC')" - is the industry standard parameter of pound-mass (lbm) of fuel per hour being burned divided by pound-force (lbf) of thrust the engine produces at that minimum point.
  • Low fan pressure ratio is the pressure ratio across the fan blade alone, without a Fan Exit Guide Vane (“FEGV”) system.
  • the low fan pressure ratio as disclosed herein according to one non- limiting embodiment is less than about 1.50. In another non- limiting embodiment the low fan pressure ratio is less than about 1.45.
  • Low corrected fan tip speed is the actual fan tip speed in ft/sec divided by an industry standard temperature correction of [(Tram °R)/ (518.7 °R)] ° '5 .
  • the "Low corrected fan tip speed”, as disclosed herein according to one non-limiting embodiment, is less than about 1150 ft/second.
  • the example gas turbine engine includes the fan 42 that comprises in one non-limiting embodiment less than about 26 fan blades. In another non-limiting embodiment, the fan section 22 includes less than about 20 fan blades. Moreover, in one disclosed embodiment the low pressure turbine 46 includes no more than about 6 turbine rotors schematically indicated at 34. In another non-limiting example embodiment the low pressure turbine 46 includes about 3 turbine rotors. A ratio between the number of fan blades 42 and the number of low pressure turbine rotors is between about 3.3 and about 8.6. The example low pressure turbine 46 provides the driving power to rotate the fan section 22 and therefore the relationship between the number of turbine rotors 34 in the low pressure turbine 46 and the number of blades 42 in the fan section 22 disclose an example gas turbine engine 20 with increased power transfer efficiency.
  • a center of gravity indicated at 62 is transferred radially to an engine centerline disposed along the axis A.
  • the center of gravity 62 is positioned a first distance XCG between forward and aft ends of the engine 20.
  • the forward end is a leading edge 66 of the fan blades 42 and the aft end is a trailing edge 68 of an aft most rotating turbine blade 70.
  • a total length XL is defined between the leading edge 66 and the trailing edge 68.
  • the position of the center of gravity 62 along the axis A is disposed at the first distance XCG and influences the configuration for supporting engines on an airframe and is therefore of concern to engine manufactures and aircraft designers.
  • a mounting structure schematically indicated at 64 supports the example gas turbine engine 20 on an airframe (not shown). It should be understood, that the location of the mounting structure 64 may vary for each engine application and such variations are within the contemplation of this disclosure.
  • the example gas turbine engine 20 includes the geared architecture 48 for driving the fan section 22 at a speed different than that of a fan drive turbine.
  • the fan drive turbine is the low pressure turbine 46.
  • the selection of materials and component configurations define the overall weight of the engine 20 along with the distribution of that weight to determine the location of the center of gravity 62.
  • Structures such as the fan section 22 and the geared architecture 48 along with the compressor section 24 and turbine section 28 combine to define not only the overall weight of the engine 20, but also the distribution of that weight that determines the location of the center of gravity 62.
  • the geared architecture 48 is a structure located forward in the engine 20 along with the fan section 22 and therefore material selection and structure configurations influence the location of the center of gravity 62. Moreover, many structures within the gas turbine engine structures factor into and determine in part the positioning and location of the center of gravity 62.
  • components of the geared architecture 48 such as for example journal bearings, lubrication jets, and a gutter around epicyclic components are selected to provide a weight reduction and an axial size reduction that define an overall weight of the geared architecture 48 and thereby factor into the definition of the center of gravity 62.
  • the weight of the geared architecture is stated as a density to relate the overall size or volume of the geared architecture to a weight.
  • the example geared architecture is a gearbox having a density between about 0.22 lbs/in 3 and about 0.30 lbs/in 3 . The part density of the gearbox is one consideration that influences the position of the center of gravity 62.
  • the fan section 22 is the one of the forward most components of the example engine and therefore also has a significant influence on the final location of the center of gravity 62 in a completed engine 22. Accordingly, selection of parts making up the fan section 22 is considered not only in view of propulsive efficiencies, but also in regard to the overall structure of the gas turbine engine 20.
  • the fan section 22 includes a rotor 72 that supports the plurality of fan blades 42 for rotation about the axis A.
  • the rotor 72 arrangement that folds radially back around the bearings 38 provides a weight benefit and thereby provides another means for modifying and positioning the engine center of gravity 62.
  • the rotor bearing 38 can include tapered roller bearings that further provide a beneficial impact on the structure of the fan section without adding additional weight to further influence the location of the center of gravity 62.
  • a density of the fan rotor can be further selected to utilize light weight structures that further correspond and effect the location of the center of gravity 62.
  • the fan section 22, including the rotor 72 and the plurality of fan blades combine to provide a density within a range of between about 0.0094 lbs/in 3 and about 0.01540 lbs/in 3 .
  • a fan containment case 16 is required to contain the blades 42 and is fabricated from a composite material to reduce weight and is a selection that determines the location of the center of gravity 62.
  • the example geared turbofan engine has an internal moment arm that is a measure of the location of the engine center of gravity 62.
  • the example internal moment arm is within a range of between about 40% and 70% of the length of the engine 20 according to the following relationship.
  • the internal moment arm MA is defined as: [0055]
  • the position of the center of gravity 62 is related as a moment arm according to the above equation and is within a range of between about 35% and about 75% of the total length XL between the leading edge 66 of the fan blades 42 and the trailing edge 68 of the aft most rotating turbine blade 70.
  • the example moment arm MA for the example engine 20 is within a range of between about 40% and about 70% of the total length XL between the leading edge 66 of the fan blades 42 and the trailing edge 68 of the aft most rotating turbine blade 70.
  • the disclosed moment arms MA represents the location of the center of gravity 62 as a percentage of the engine length XL ( Figure 1) measured between the leading edge 66 of the fan blade 72 and the trailing edge 68 of the last rotating turbine blade 70 in the low pressure turbine 46.
  • a hoist point 74 located on the engine 20 will be axially located at the center of gravity 62 when the engine center line or axis A is normal to a vertical line 76 extending through the hoist point 74 and intersecting the axis A. Accordingly, it this example the center of gravity 62 may in one example be determined as that point where the vertical line 76 intersects the engine center line or axis A at a right angle when supported at a single hoist point 74.
  • the example engine center of gravity 62 is considered along the engine centerline or axis A although the actual center of gravity 62 may be slightly skewed from the axis A due to locally mounted accessory components.
  • the disclosed center of gravity 62 includes fluids contained within operating systems of the turbofan engine 20.
  • the example center of gravity does not include some structures such as for example typical tubes, brackets and harness such as those coming from the airframe which would have almost no effect on the location of the center of gravity 62.
  • the example center of gravity 62 is determined for a bare engine only that does not include engine mounts, a fan cowl, a thrust reverser, an inlet, nozzle or plug.
  • the example center of gravity 62 is not determined including weights of structures comprising the gas turbine engine not including engine mounting structures, engine cowling structures and nacelle structures.
  • the center of gravity 62 can be located in a structurally desirable location to increase propulsive efficiencies and reduce mounting structure requirements.

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  • Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Structures Of Non-Positive Displacement Pumps (AREA)

Abstract

Cette invention concerne une turbine à gaz, comprenant une pluralité d'aubes de soufflante tournant autour d'un axe, chacune desdites aubes de soufflante comprenant un bord d'attaque. Ladite turbine à gaz comprend en outre une section de turbine comprenant une aube de turbine disposée le plus en arrière présentant un bord de fuite et une architecture engrenée entraînée par la section de turbine pour entraîner en rotation la pluralité d'aubes de turbine autour de l'axe. Un centre de gravité de ladite turbine à gaz est situé à une première distance axiale du bord de fuite de l'aube de turbine la plus en arrière, qui va d'environ 35 à environ 75 % d'une longueur totale entre le bord d'attaque de la pluralité d'aubes de turbine et le bord de fluide de l'aube de turbine la plus en arrière.
PCT/US2013/062104 2012-10-01 2013-09-27 Turbine à gaz à bras de levier avant Ceased WO2014099085A2 (fr)

Priority Applications (5)

Application Number Priority Date Filing Date Title
US14/432,289 US11635025B2 (en) 2012-10-01 2013-09-27 Gas turbine engine with forward moment arm
CA2884976A CA2884976C (fr) 2012-10-01 2013-09-27 Turbine a gaz a bras de levier avant
BR112015006820-0A BR112015006820B1 (pt) 2012-10-01 2013-09-27 Motores de turbina a gás e de ventilador turbo, e, método para montar um motor de turbina a gás
JP2015533317A JP6027250B2 (ja) 2012-10-01 2013-09-27 前方モーメントアームを有するガスタービンエンジン
EP13865250.8A EP2904210A4 (fr) 2012-10-01 2013-09-27 Turbine à gaz à bras de levier avant

Applications Claiming Priority (4)

Application Number Priority Date Filing Date Title
US201261708240P 2012-10-01 2012-10-01
US61/708,240 2012-10-01
US201361789275P 2013-03-15 2013-03-15
US61/789,275 2013-03-15

Publications (2)

Publication Number Publication Date
WO2014099085A2 true WO2014099085A2 (fr) 2014-06-26
WO2014099085A3 WO2014099085A3 (fr) 2014-09-18

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PCT/US2013/062104 Ceased WO2014099085A2 (fr) 2012-10-01 2013-09-27 Turbine à gaz à bras de levier avant

Country Status (5)

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EP (1) EP2904210A4 (fr)
JP (1) JP6027250B2 (fr)
BR (1) BR112015006820B1 (fr)
CA (1) CA2884976C (fr)
WO (1) WO2014099085A2 (fr)

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Cited By (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US20250122848A1 (en) * 2022-06-22 2025-04-17 General Electric Company Gearbox assembly with lubricant extraction volume ratio

Also Published As

Publication number Publication date
JP2015534622A (ja) 2015-12-03
CA2884976A1 (fr) 2014-06-26
JP6027250B2 (ja) 2016-11-16
BR112015006820A2 (pt) 2017-07-04
EP2904210A4 (fr) 2015-12-02
BR112015006820B1 (pt) 2022-12-06
CA2884976C (fr) 2018-03-06
WO2014099085A3 (fr) 2014-09-18
EP2904210A2 (fr) 2015-08-12

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