EP3088674A1 - Pale de rotor et turbine à gaz associée - Google Patents

Pale de rotor et turbine à gaz associée Download PDF

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Publication number
EP3088674A1
EP3088674A1 EP16166812.4A EP16166812A EP3088674A1 EP 3088674 A1 EP3088674 A1 EP 3088674A1 EP 16166812 A EP16166812 A EP 16166812A EP 3088674 A1 EP3088674 A1 EP 3088674A1
Authority
EP
European Patent Office
Prior art keywords
side wall
tip
airfoil
suction side
rotor blade
Prior art date
Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
Granted
Application number
EP16166812.4A
Other languages
German (de)
English (en)
Other versions
EP3088674B1 (fr
Inventor
Jeffrey Clarence JONES
Xiuzhang Zhang
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.)
General Electric Technology GmbH
Original Assignee
General Electric Co
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 General Electric Co filed Critical General Electric Co
Publication of EP3088674A1 publication Critical patent/EP3088674A1/fr
Application granted granted Critical
Publication of EP3088674B1 publication Critical patent/EP3088674B1/fr
Active legal-status Critical Current
Anticipated expiration legal-status Critical

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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
    • F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/12—Blades
    • F01D5/14—Form or construction
    • F01D5/141—Shape, i.e. outer, aerodynamic form
    • 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
    • F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/12—Blades
    • F01D5/14—Form or construction
    • F01D5/20—Specially-shaped blade tips to seal space between tips and stator
    • 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
    • F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/12—Blades
    • F01D5/14—Form or construction
    • F01D5/18—Hollow blades, i.e. blades with cooling or heating channels or cavities; Heating, heat-insulating or cooling means on blades
    • F01D5/187—Convection cooling
    • F01D5/188—Convection cooling with an insert in the blade cavity to guide the cooling fluid, e.g. forming a separation wall
    • F01D5/189—Convection cooling with an insert in the blade cavity to guide the cooling fluid, e.g. forming a separation wall the insert having a tubular cross-section, e.g. airfoil shape
    • 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
    • F05D2250/00—Geometry
    • F05D2250/30—Arrangement of components
    • F05D2250/31—Arrangement of components according to the direction of their main axis or their axis of rotation
    • F05D2250/314—Arrangement of components according to the direction of their main axis or their axis of rotation the axes being inclined in relation to each other
    • 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
    • F05D2250/00—Geometry
    • F05D2250/30—Arrangement of components
    • F05D2250/32—Arrangement of components according to their shape
    • F05D2250/324—Arrangement of components according to their shape divergent
    • 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
    • F05D2250/00—Geometry
    • F05D2250/70—Shape
    • F05D2250/71—Shape curved
    • F05D2250/713—Shape curved inflexed

Definitions

  • the present invention generally relates to a rotor blade for a turbine. More particularly, this invention involves a rotor blade having a flared tip configured for cooling a trailing edge portion of the rotor blade.
  • an air-ingesting turbo machine e.g., a gas turbine
  • air is pressurized by a compressor and then mixed with fuel and ignited within an annular array of combustors to generate combustion gases.
  • the hot gases are routed through a liner and into a hot gas path defined within a turbine section of the turbo machine.
  • Kinetic energy is extracted from the combustion gases via one or more rows of turbine rotor blades that are connected to a rotor shaft. The extracted kinetic energy causes the rotor shaft to rotate, thus producing work.
  • the turbine rotor blades or blades generally operate in extremely high temperature environments.
  • the blades typically include various internal cooling passages or cavities.
  • a cooling medium such as compressed air is routed through the internal cooling passages.
  • a portion of the cooling medium may be routed out of the internal cooling passages through various cooling holes defined along the blade surface, thereby reducing high surface temperatures.
  • An area that is generally challenging to cool effectively via the cooling medium is a blade tip portion of the turbine rotor blade, more particularly a trailing edge region of the blade tip.
  • the blade tip is generally defined at a radial extremity of the turbine rotor blade and is positioned radially inward from a turbine shroud that circumscribes the row of blades.
  • the turbine shroud defines a radially outward boundary of the hot gas path. The proximity of the blade tip to the turbine shroud makes the blade tip difficult to cool. The contiguity of the shroud and the blade tip minimizes the leakage of hot operating fluid past the tip which correspondingly improves turbine efficiency.
  • a tip cavity formed by a recessed tip cap and a pressure side wall and a suction side wall provides a means for achieving minimal tip clearance while at the same time assuring adequate blade tip cooling.
  • the pressure side wall and the suction side wall extend radially outwardly from the tip cap. At least a portion of at least one of the suction side wall and the pressure side wall is flared or inclined outward with respect to a radial centerline of the blade.
  • the pressure side wall intersects with the suction side wall at a leading edge portion of the blade. However, the pressure side wall does not intersect with the suction side wall at the trailing edge, thus forming an opening therebetween. This configuration is generally due to the lack of an appropriate wall thickness of the blade along the trialing edge.
  • the cooling medium is exhausted from the internal passages through holes in the tip cap into the tip cavity, thus effectively cooling the pressure and suction side walls as well as the tip cap surface.
  • the rotor blade includes an airfoil having a pressure side wall and a suction side wall that are connected at leading and trailing edges of the airfoil.
  • a blade tip defines a radially outer surface of the airfoil.
  • the rotor blade also includes an internal cavity for receiving a cooling medium.
  • a tip cavity that is in fluid communication with the internal cavity is at least partially defined by a tip cap.
  • the tip cavity is recessed radially inwardly from the radially outer surface and surrounded by the pressure and suction side walls.
  • a portion of at least one of the suction side wall or the pressure side wall that defines the tip cavity extends obliquely outwardly from the tip cavity.
  • a plurality of slots is defined in at least one of the suction side wall or the pressure side wall along the radially outer surface proximate to the trailing edge of the airfoil.
  • the gas turbine includes, in serial flow order, a compressor section, a combustion section and a turbine section.
  • the turbine section includes a rotor shaft and a plurality of rotor blades that are coupled to the rotor blade.
  • Each rotor blade includes an airfoil having a pressure side wall and a suction side wall that are connected at leading and trailing edges of the airfoil.
  • a blade tip defines a radially outer surface of the airfoil.
  • the rotor blade also includes an internal cavity for receiving a cooling medium.
  • a tip cavity that is in fluid communication with the internal cavity is at least partially defined by a tip cap.
  • the tip cavity is recessed radially inwardly from the radially outer surface and surrounded by the pressure and suction side walls.
  • a portion of at least one of the suction side wall or the pressure side wall that defines the tip cavity extends obliquely outwardly from the tip cavity.
  • a plurality of slots is defined in at least one of the suction side wall or the pressure side wall along the radially outer surface proximate to the trailing edge of the airfoil.
  • radially refers to the relative direction that is substantially perpendicular to an axial centerline of a particular component and/or substantially perpendicular to an axial centerline of the turbomachine
  • axially refers to the relative direction that is substantially parallel and/or coaxially aligned to an axial centerline of a particular component and/or to an axial centerline of the turbomachine.
  • FIG. 1 illustrates a schematic diagram of one embodiment of a gas turbine 10.
  • the gas turbine 10 generally includes an inlet section 12, a compressor section 14 disposed downstream of the inlet section 12, a plurality of combustors (not shown) within a combustor section 16 disposed downstream of the compressor section 14, a turbine section 18 disposed downstream of the combustor section 16 and an exhaust section 20 disposed downstream of the turbine section 18. Additionally, the gas turbine 10 may include one or more shafts 22 coupled between the compressor section 14 and the turbine section 18.
  • the turbine section 18 may generally include a rotor shaft 24 having a plurality of rotor disks 26 (one of which is shown) and a plurality of rotor blades 28 extending radially outwardly from and being interconnected to the rotor disk 26. Each rotor disk 26 in turn, may be coupled to a portion of the rotor shaft 24 that extends through the turbine section 18.
  • the turbine section 18 further includes an outer casing 30 that circumferentially surrounds the rotor shaft 24 and the rotor blades 28, thereby at least partially defining a hot gas path 32 through the turbine section 18.
  • a working fluid such as air flows through the inlet section 12 and into the compressor section 14 where the air is progressively compressed, thus providing pressurized air to the combustors of the combustion section 16.
  • the pressurized air is mixed with fuel and burned within each combustor to produce combustion gases 34.
  • the combustion gases 34 flow through the hot gas path 32 from the combustor section 16 into the turbine section 18, wherein energy (kinetic and/or thermal) is transferred from the combustion gases 34 to the rotor blades 28, thus causing the rotor shaft 24 to rotate.
  • the mechanical rotational energy may then be used to power the compressor section 14 and/or to generate electricity.
  • the combustion gases 34 exiting the turbine section 18 may then be exhausted from the gas turbine 10 via the exhaust section 20.
  • FIG. 2 is a perspective view of an exemplary rotor blade 100 as may incorporate one or more embodiments of the present invention and as may be incorporated into the turbine section 18 of the gas turbine 10 in place of rotor blade 28 as shown in FIG. 1 .
  • the rotor blade 100 generally includes a mounting or shank portion 102 having a mounting body 104, and an airfoil 106 that extends in span outwardly in a radial direction 108 from a platform portion 110 of the rotor blade 100.
  • the platform 110 generally serves as a radially inward flow boundary for the combustion gases 34 flowing through the hot gas path 32 of the turbine section 18 ( FIG. 1 ).
  • FIG. 1 As shown in FIG.
  • the mounting body 104 of the mounting or shank portion 102 may extend radially inwardly from the platform 110 and may include a root structure, such as a dovetail, configured to interconnect or secure the rotor blade 100 to the rotor disk 26 ( FIG. 1 ).
  • the airfoil 106 includes an outer surface 112 that surrounds the airfoil 106.
  • the outer surface 112 is at least partially defined by a pressure side wall 114 and an opposing suction side wall 116.
  • the pressure side wall 114 and the suction side wall 116 extend substantially radially outwardly from the platform 110 in span from a root 118 of the airfoil 106 to a blade tip or tip 120 of the airfoil 106.
  • the root 118 of the airfoil 106 may be defined at an intersection between the airfoil 106 and the platform 110.
  • the blade tip 120 is disposed radially opposite the root 118. As such, a radially outer surface 122 of the blade the tip 120 may generally define the radially outermost portion of the rotor blade 100.
  • the pressure side wall 114 and the suction side wall 116 are joined together or interconnected at a leading edge 124 of the airfoil 106 which is oriented into the flow of combustion gases 34.
  • the pressure side wall 114 and the suction side wall 116 are also joined together or interconnected at a trailing edge 126 of the airfoil 106 which is spaced downstream from the leading edge 124.
  • the pressure side wall 114 and the suction side wall 116 are continuous about the trailing edge 126.
  • the pressure side wall 114 is generally concave and the suction side wall 116 is generally convex.
  • chord of the airfoil 106 is the length of a straight line connecting the leading edge 114 and the trailing edge 116 and the direction from the leading edge 114 to the trailing edge 116 is typically described as the chordwise direction.
  • a chordwise line bisecting the pressure side wall 114 and the suction side wall 116 is typically referred to as the mean-line or camber-line 128 of the airfoil 106.
  • a cooling medium such as a relatively cool compressed air bled from the compressor section 14 ( FIG. 1 ) of the gas turbine engine 10 which is suitably channeled through the mounting or shank portion 102 of the rotor blade 100 and into an internal cavity or passage 132 that is at least partially defined within the airfoil 106 between the pressure side wall 114 and the suction side wall 116.
  • the internal cavity 132 may take any conventional form and is typically in the form of a serpentine passage.
  • the cooling medium 130 enters the internal cavity 132 from the mounting or shank portion 102 and passes through the internal cavity 132 for suitably cooling the airfoil 106 from the heating effect of the combustion gases 34 flowing over the outer surface 112 thereof.
  • Film cooling holes may be disposed on the pressure side wall 114 and/or the suction side wall 116 for conventionally film cooling the outer surface 112 of the airfoil 106.
  • a tip cavity or plenum 134 is formed at or within the blade tip 120.
  • the tip cavity 134 is at least partially formed by a tip cap 136.
  • the tip cap 136 is recessed radially inwardly from the blade tip 120 and/or the outer surface 122 of the blade tip 120 and forms a floor portion of the tip cavity 134.
  • the tip cap 136 is surrounded continuously by the pressure side wall 114 and the suction side wall 116.
  • the tip cap 136 is connected to and/or forms a seal against an inner surface or side 138 of the pressure side wall 114 and an inner surface or side 140 of the suction side wall 116 along a periphery 142 of the tip cap 136 between the leading and trailing edges 124, 126 of the airfoil 106.
  • the tip cap 136 further includes a plurality of holes or apertures 144 that extend through a top surface or side 146 of the tip cap 136 and that provide for fluid communication between the internal cavity 132 and the tip cavity 134.
  • FIG. 3 provides a perspective view of a portion the airfoil 106 which includes the blade tip 120 according to at least one embodiment of the present invention.
  • FIG. 4 provides a cross sectioned top view of a portion of the airfoil 106 taken along section lines 4-4 as shown in FIG. 3 , according to at least one embodiment of the present invention.
  • FIG. 5 provides a cross sectioned side view of a portion of the airfoil 106 taken along section lines 5-5 as shown in FIG. 3 , according to at least one embodiment of the present invention.
  • a portion of at least one of the suction side wall 116 or the pressure side wall 114 that defines the tip cavity 134 extends obliquely outwardly from the tip cavity 134 and/or the top surface 146 of the tip cap 136 with respect to radial direction 108 and/or with respect to the outer surface 112 of the airfoil 106.
  • Radial direction 108 may be substantially perpendicular to the top surface 146 of the tip cap 136.
  • a portion of the suction side wall 116 that defines the tip cavity 134 and a portion of the pressure side wall 114 that defines the tip cavity 134 extends obliquely outwardly from the tip cavity 134 with respect to radial direction 108 and/or with respect to the outer surface 112 of the airfoil 106.
  • a portion of the suction side wall 116 that defines the tip cavity 134 extends obliquely outwardly from the tip cavity 134 with respect to radial direction 108 and/or with respect to the outer surface 112 of the airfoil 106.
  • a portion of the pressure side wall 114 that defines the tip cavity 134 extends obliquely outwardly from the tip cavity 134 with respect to radial direction 108 and/or with respect to the outer surface 112 of the airfoil 106.
  • a portion of the inner surface or side 140 of the suction side wall 116 that defines the tip cavity 134 may extend obliquely outwardly from the tip cavity 134 with respect to radial direction 108, thus increasing an overall volume of the tip cavity 134.
  • a portion of the inner surface or side 138 of the pressure side wall 114 that defines the tip cavity 134 may extend obliquely outwardly from the tip cavity 134 with respect to radial direction 108, thus increasing an overall volume of the tip cavity 134.
  • the airfoil 106 includes a plurality of slots 148 defined by or within at least one of the suction side wall 116 or the pressure side wall 114 along the radially outer surface 122 and positioned proximate to the trailing edge 126 of the airfoil 106. As shown in FIGS. 3 and 4 , the pressure side wall 114 and the suction side wall 116 maintain continuity across the trailing edge 126. Although the plurality of slots 148 are shown in FIGS.
  • the plurality of slots 148 may occur only along the suction side wall 116 or occur only along the pressure side wall 114 or may occur along both the pressure side and the suction side walls 114, 116 as shown.
  • the plurality of slots 148 occurs only along the suction side wall 116. In another embodiment, the plurality of slots 148 occurs only along the pressure side wall 114. In one embodiment, as shown in FIG. 4 , the plurality of slots 148 includes a first slot 150 defined in the pressure side wall 114 and a second slot 152 defined within the suction side wall along the radially outer surface proximate to the trailing edge 126 of the airfoil 106. In particular embodiments, the plurality of slots 148 is equally or non-equally distributed on both the pressure and suction side walls 114, 116.
  • one or more slots 148 of the plurality of slots 148 extend through the radially outer surface 122 of the airfoil 106 towards the top surface 146 of the tip cap 136.
  • one or more of the slots 148 may be angled towards the trialing edge 126 as it extends through the inner surface 138 of the pressure side wall 114 or the inner surface 140 of the suction side wall 116 and the outer surface 112 of the airfoil 106.
  • at least one slot 148 of the plurality of slots 148 extends radially into and/or at least partially through the top surface 146 of the tip cap 136.
  • FIG. 6 provides a perspective view of a portion the airfoil 106 which includes the blade tip 120 according to at least one embodiment of the present invention.
  • FIG. 7 provides a cross sectioned side view of a portion of the airfoil 106 taken along section lines 7-7 as shown in FIG. 6 , according to at least one embodiment of the present invention.
  • a portion 154 of the top surface 146 of the tip cap 136 that is proximate to the trailing edge 126 is stepped radially inwardly. Stepped portion 154 may be inclined along the camber line 128 ( FIG. 2 ) or otherwise contoured to facilitate or enhance cooling effectiveness.
  • one or more of the slots 148 of the plurality of slots 148 may be tapered and/or non-linear to enable flow of the cooling medium from the tip cavity 134 adjacent to the trailing edge 126.
  • at least one slot 148 of the plurality of slots 148 may taper inwardly from the outer radial surface 122 of the blade tip 120 towards the top surface 146 of the tip cap 136.
  • at least one slot 148 of the plurality of slots 148 may include a curved or widened inlet 156 along inner surfaces 138, 140.
  • at least one slot 148 may also include a widened or diffusing outlet 158 along the outer surface 112.
  • at least one slot may include a narrowing region 160 defined between the inlet 156 and outlet 158.
  • At least one aperture 144 of the plurality of apertures 144 is positioned proximate to the trialing edge 126.
  • at least one aperture 148 of the plurality of apertures 144 is angled aft towards the trailing edge 126 of the airfoil 106 with respect to radial direction 108.
  • at least one aperture 144 of the plurality of apertures 144 is defined between adjacent slots 148 of the plurality of slots 148.
  • at least one aperture 144 may be disposed upstream of at least one slot 148.
  • one or more holes 162 are defined along the trailing edge 126 of the airfoil 106 radially below the tip cap 136.
  • the one or more holes 162 may be in fluid communication with the internal cavity 132, thus providing additional cooling along the trailing edge 126 of the airfoil 106.

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  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Physics & Mathematics (AREA)
  • Fluid Mechanics (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
EP16166812.4A 2015-04-29 2016-04-25 Pale de rotor et turbine à gaz associée Active EP3088674B1 (fr)

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
US14/699,308 US20160319672A1 (en) 2015-04-29 2015-04-29 Rotor blade having a flared tip

Publications (2)

Publication Number Publication Date
EP3088674A1 true EP3088674A1 (fr) 2016-11-02
EP3088674B1 EP3088674B1 (fr) 2024-05-29

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Family Applications (1)

Application Number Title Priority Date Filing Date
EP16166812.4A Active EP3088674B1 (fr) 2015-04-29 2016-04-25 Pale de rotor et turbine à gaz associée

Country Status (4)

Country Link
US (1) US20160319672A1 (fr)
EP (1) EP3088674B1 (fr)
JP (1) JP6824623B2 (fr)
CN (1) CN106089313B (fr)

Cited By (1)

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Publication number Priority date Publication date Assignee Title
EP3138997A1 (fr) * 2015-09-02 2017-03-08 General Electric Company Configurations de pointes d'aube de rotor de turbine

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US10677066B2 (en) 2015-11-23 2020-06-09 United Technologies Corporation Turbine blade with airfoil tip vortex control
US20170145827A1 (en) * 2015-11-23 2017-05-25 United Technologies Corporation Turbine blade with airfoil tip vortex control
WO2017168646A1 (fr) * 2016-03-30 2017-10-05 三菱重工業株式会社 Turbocompresseur à géométrie variable
US10443405B2 (en) * 2017-05-10 2019-10-15 General Electric Company Rotor blade tip
CN107559048B (zh) * 2017-09-22 2024-01-30 哈尔滨汽轮机厂有限责任公司 一种用于中低热值重型燃气轮机发动机的转子叶片
KR102021139B1 (ko) * 2018-04-04 2019-10-18 두산중공업 주식회사 스퀼러 팁을 구비한 터빈 블레이드
US10787932B2 (en) * 2018-07-13 2020-09-29 Honeywell International Inc. Turbine blade with dust tolerant cooling system
CN110863862B (zh) * 2019-12-05 2022-12-06 中国航发四川燃气涡轮研究院 叶尖结构和涡轮
US11225874B2 (en) * 2019-12-20 2022-01-18 Raytheon Technologies Corporation Turbine engine rotor blade with castellated tip surface
US12123319B2 (en) 2020-12-30 2024-10-22 Ge Infrastructure Technology Llc Cooling circuit having a bypass conduit for a turbomachine component

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US20040179940A1 (en) * 2003-03-12 2004-09-16 Florida Turbine Technologies, Inc. Multi-metered film cooled blade tip
US8801377B1 (en) * 2011-08-25 2014-08-12 Florida Turbine Technologies, Inc. Turbine blade with tip cooling and sealing
US20140047842A1 (en) * 2012-08-15 2014-02-20 Wieslaw A. Chlus Suction side turbine blade tip cooling

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EP3138997A1 (fr) * 2015-09-02 2017-03-08 General Electric Company Configurations de pointes d'aube de rotor de turbine

Also Published As

Publication number Publication date
CN106089313B (zh) 2020-12-01
JP6824623B2 (ja) 2021-02-03
EP3088674B1 (fr) 2024-05-29
JP2016211545A (ja) 2016-12-15
CN106089313A (zh) 2016-11-09
US20160319672A1 (en) 2016-11-03

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