US20150345301A1 - Rotor blade cooling flow - Google Patents

Rotor blade cooling flow Download PDF

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Publication number
US20150345301A1
US20150345301A1 US14/289,908 US201414289908A US2015345301A1 US 20150345301 A1 US20150345301 A1 US 20150345301A1 US 201414289908 A US201414289908 A US 201414289908A US 2015345301 A1 US2015345301 A1 US 2015345301A1
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US
United States
Prior art keywords
tip
side portion
coolant
floor
pocket
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
US14/289,908
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English (en)
Inventor
Xiuzhang James Zhang
James William Vehr
Haiping Wang
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 Co
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
Priority to US14/289,908 priority Critical patent/US20150345301A1/en
Assigned to GENERAL ELECTRIC COMPANY reassignment GENERAL ELECTRIC COMPANY ASSIGNMENT OF ASSIGNORS INTEREST (SEE DOCUMENT FOR DETAILS). Assignors: VEHR, JAMES WILLIAM, WANG, Haiping, ZHANG, XIUZHANG JAMES
Priority to DE102015107844.9A priority patent/DE102015107844A1/de
Priority to JP2015103298A priority patent/JP2015224634A/ja
Priority to CH00756/15A priority patent/CH709650A2/de
Priority to CN201510285010.9A priority patent/CN105317468A/zh
Publication of US20150345301A1 publication Critical patent/US20150345301A1/en
Abandoned legal-status Critical Current

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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
    • F01D5/20—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
    • 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

  • the present invention generally relates to a rotor blade for a turbine. More particularly, this invention involves a rotor blade having a tip configured for promoting coolant flow through the rotor blade.
  • 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 hot gases of combustion.
  • the hot gases flow from each combustor through a transition piece for flow along an annular hot gas path.
  • Turbine stages are typically disposed along the hot gas path such that the hot gases flow across first-stage nozzles and rotor blades and across the nozzles and rotor blades of follow-on turbine stages.
  • the rotor blades may be secured to a plurality of rotor disks which are coupled to a turbine rotor shaft, with each rotor disk being mounted to the rotor shaft.
  • a rotor blade generally includes an airfoil that extends radially outwardly from a substantially planar platform and a mounting portion that extends radially inwardly from the platform for securing the rotor blade to one of the rotor disks.
  • a tip of the airfoil is typically spaced radially inwardly from a stationary shroud or seal of the turbine such that a small clearance gap is defined between the tip and the shroud.
  • a plurality of cooling passages is defined within the airfoil for routing a coolant such as compressed air through the airfoil. In particular configurations, a plurality of coolant outlets are defined along the tip for routing the coolant out of the cooling passages at the tip.
  • the flow of the coolant through the cooling passages is primarily driven by a pressure difference defined between a supply pressure of the coolant and a static pressure which is typically defined at the tip of the airfoil at or just downstream from the coolant outlets. If the supply pressure is too low, for example, due to aerodynamic loading optimization, a decrease in operating speed and/or a change in turbine load requirements, a lower or reduced static pressure is required in order to meet cooling flow needs. Therefore, an improved rotor blade tip design which provides a lower or reduced static pressure at the tip to increase or enhance coolant flow through the airfoil would be useful.
  • One embodiment of the present invention is a rotor blade having an airfoil.
  • the airfoil includes pressure and suction side walls which extend radially outwardly from a platform in span from a root to a tip and in chord between a leading edge and a trialing edge.
  • the tip includes a tip floor and a plurality of coolant outlets disposed along the tip floor.
  • the tip further includes a tip rail that extends radially outwardly from the tip floor.
  • the tip rail has a pressure side portion and a suction side portion which are joined at the leading and trailing edges.
  • a plurality of cooling passages is circumscribed within the airfoil for routing a coolant therethrough. Each or at least some of the cooling passages is/are in fluid communication with one or more of the coolant outlets.
  • a baffle extends radially outwardly from and transversely across the tip floor from the pressure side portion to the suction side portion so as to define a first tip pocket and a second tip pocket.
  • a slot is disposed along the suction side portion of the tip rail and provides for fluid communication out of one of the first or second tip pockets.
  • the present invention is a system for promoting coolant flow through a rotor blade.
  • the system includes a coolant source for supplying a pressurized coolant to a cooling passage inlet formed along the rotor blade.
  • the rotor blade comprises a mounting portion which includes a mounting body.
  • the mounting body is interconnectable with a rotor shaft.
  • At least one of the cooling passage inlets is formed by the mounting body.
  • An airfoil extends radially outwardly from the mounting portion and includes pressure and suction side walls which extend radially outwardly from a platform in span from a root to a tip and in chord between a leading edge and a trialing edge.
  • the tip includes a tip floor and a plurality of coolant outlets disposed along the tip floor.
  • the tip further includes a tip rail that extends radially outwardly from the tip floor.
  • the tip rail includes a pressure side portion and a suction side portion which are joined at the leading and trailing edges.
  • a plurality of cooling passages is circumscribed within the airfoil for routing a coolant therethrough. Each cooling passage is in fluid communication with one or more of the coolant inlets.
  • a baffle extends radially outwardly from and transversely across the tip floor from the pressure side portion to the suction side portion to define a first tip pocket and a second tip pocket.
  • a slot is disposed along the suction side portion of the tip rail and provides for fluid communication out of one of the first or second tip pockets.
  • the gas turbine includes a compressor, a combustor disposed downstream from the compressor and a turbine disposed downstream from the combustor.
  • the turbine includes a rotor shaft that extends axially through the turbine.
  • An outer casing circumferentially surrounds the rotor shaft to define a hot gas path therebetween.
  • a plurality of rotor blades is interconnected to the rotor shaft, which together, define a stage of rotor blades.
  • Each rotor blade comprises a mounting portion which includes a mounting body.
  • the mounting body is interconnectable with a rotor shaft and at least one of the cooling passage inlets is formed in the mounting body.
  • the rotor blade further includes an airfoil that is coupled to the mounting portion and that includes pressure and suction side walls which extend radially outwardly from a platform in span from a root to a tip and in chord between a leading edge and a trialing edge.
  • the tip includes a tip floor and a plurality of coolant outlets disposed along the tip floor.
  • the tip further includes a tip rail that extends radially outwardly from the tip floor.
  • the tip rail includes a pressure side portion and a suction side portion which are joined at the leading and trailing edges.
  • a plurality of cooling passages is circumscribed within the airfoil for routing a coolant through the airfoil. Each cooling passage is in fluid communication with one or more of the coolant inlets.
  • a baffle extends radially outwardly from and transversely across the tip floor from the pressure side portion to the suction side portion to define a first tip pocket and a second tip pocket. At least one coolant outlet is disposed along the tip floor within the first tip pocket and at least one coolant outlet is disposed along the tip floor within the second tip pocket.
  • a slot is disposed along the suction side portion of the tip rail and provides for fluid communication out of one of the first or second tip pockets.
  • FIG. 1 illustrates a functional diagram of an exemplary gas turbine as may incorporate at least one embodiment of the present invention
  • FIG. 2 is a perspective view of an exemplary rotor blade as may incorporate various embodiments of the present disclosure
  • FIG. 3 is an enlarged perspective view of a tip of an exemplary rotor blade, according to at least one embodiment of the invention.
  • FIG. 4 is an enlarged top view of the exemplary rotor blade tip as shown in FIG. 3 ;
  • FIG. 5 is an enlarged perspective view of an exemplary rotor blade, according to at least one embodiment of the invention.
  • FIG. 6 is an enlarged perspective view of an exemplary rotor blade, according to at least one embodiment of the invention.
  • FIG. 7 is an enlarged perspective view of an exemplary rotor blade, according to at least one embodiment of the invention.
  • FIG. 8 is an enlarged perspective view of an exemplary rotor blade, according to at least one embodiment of the invention.
  • upstream and downstream refer to the relative direction with respect to fluid flow in a fluid pathway.
  • upstream refers to the direction from which the fluid flows
  • downstream refers to the direction to which the fluid flows.
  • radially refers to the relative direction that is substantially perpendicular to an axial centerline of a particular component
  • axially refers to the relative direction that is substantially parallel and/or coaxially aligned to an axial centerline of a particular component.
  • 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 .
  • 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 each rotor disk 26 . Each rotor disk 26 may, in turn, 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 hot gases of combustion 34 .
  • the hot gases of combustion 34 flow through the hot gas path 32 from the combustor section 16 to the turbine section 18 , wherein energy (kinetic and/or thermal) is transferred from the hot 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 generate electricity.
  • the hot gases of combustion 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 28 as may incorporate one or more embodiments of the present invention.
  • the rotor blade 28 generally includes a mounting or shank portion 36 having a mounting body 38 and an airfoil 40 that extends substantially radially outwardly from a substantially planar platform 42 .
  • the platform 42 generally serves as the radially inward boundary for the hot gases of combustion 34 flowing through the hot gas path 32 of the turbine section 18 ( FIG. 1 ).
  • FIG. 1 As shown in FIG.
  • the mounting body 38 of the mounting or shank portion 36 may extend radially inwardly from the platform 42 and may include a root structure, such as a dovetail, configured to interconnect or secure the rotor blade 28 to the rotor disk 26 ( FIG. 1 ).
  • the airfoil 40 includes a pressure side wall 44 and an opposing suction side wall 46 .
  • the pressure side wall 44 and the suction side wall 46 extend substantially radially outwardly from the platform 42 in span from a root 48 of the airfoil 40 which may be defined at an intersection between the airfoil 40 and the platform 42 , and a tip 50 of the airfoil 40 .
  • the pressure side wall 44 and suction side wall 46 extend in chord between a leading edge 52 and a trialing edge 54 of the airfoil 40 .
  • the pressure side wall 44 generally comprises an aerodynamic, concave outer surface of the airfoil 40 .
  • the suction side wall 46 may generally define an aerodynamic, convex outer surface of the airfoil 40 .
  • the tip 50 is disposed radially opposite the root. As such, the tip 50 may generally define the radially outermost portion of the rotor blade 28 and thus, may be configured to be positioned adjacent to a stationary shroud or seal (not shown) of the gas turbine 10 .
  • a plurality of cooling passages 56 (shown in dashed lines in FIG. 2 ) is circumscribed within the airfoil 40 for routing a coolant 58 through the airfoil 40 between the pressure side wall 44 and the suction side wall 46 , thus providing convective cooling thereto.
  • the coolant 58 may include a portion of the compressed air from the compressor section 14 ( FIG. 1 ) and/or steam or any other suitable fluid or gas for cooling the airfoil 40 .
  • One or more cooling passage inlets 60 are disposed along the rotor blade 28 . In one embodiment, one or more cooling passage inlets 60 are formed within, along or by the mounting body 38 .
  • the cooling passage inlets 60 are in fluid communication with at least one corresponding cooling passage 56 .
  • FIG. 3 is an enlarged perspective view of the tip 50 of the airfoil 40 as shown in FIG. 2 , according to one embodiment of the present invention.
  • FIG. 4 is an enlarged top view of the tip 50 of the airfoil 40 as shown in FIGS. 2 and 3 .
  • the tip 50 includes a tip floor 62 .
  • the tip floor 62 generally extends between the pressure and suction side walls 44 , 46 and the leading and trailing edges 52 , 54 of the airfoil 40 .
  • a plurality of coolant outlets 64 is disposed along the tip floor 62 .
  • Each cooling passage 56 ( FIG. 2 ) is in fluid communication with at least one of the coolant outlets 64 .
  • a tip rail 66 extends radially outwardly from the tip floor 64 .
  • the tip rail 66 comprises a pressure side portion 68 and a suction side portion 70 .
  • the pressure side portion 68 extends along a perimeter of the tip floor 62 and generally conforms in profile to the pressure side wall 44 .
  • the suction side portion 70 extends along the perimeter of the tip floor 62 and generally conforms in profile to the suction side wall 46 .
  • the pressure side portion 68 and the suction side portion 70 are joined and/or intersect at the leading edge 52 and at and/or proximate to the trailing edge 54 .
  • a baffle 72 extends radially outwardly from the tip floor 62 .
  • the baffle 72 extends across the tip floor 62 from the pressure side to the suction side portions 68 , 70 of the tip rail 66 .
  • the baffle 72 , the tip rail 66 and the tip floor 62 define a first tip pocket 74 and a second tip pocket 76 along the tip 50 of the airfoil 40 .
  • the first tip pocket 74 is generally defined adjacent and/or proximate to the leading edge 52 of the airfoil 40 .
  • the second tip pocket 76 generally extends from the baffle 72 towards the trailing edge 54 of the airfoil 40 .
  • at least a portion of the coolant outlets 64 are formed along the tip floor 62 within the first tip pocket 74 and at least a portion of the coolant outlets 64 are formed along the tip floor 62 within the second tip pocket 76 .
  • the hot gases 34 are directed onto the pressure side wall 44 of the airfoil 40 , thus creating a high pressure region 78 along the pressure side wall 44 of each rotor blade 28 .
  • a reduced or low pressure (with respect to the high pressure region) region 80 develops along the suction side wall 46 .
  • the coolant 58 is supplied from a coolant source, such as the compressor section 14 ( FIG. 1 ) to the cooling passages 56 through the coolant passage inlets 60 at various supply pressures which generally relate to the various operating modes of the gas turbine 10 .
  • the flow of the coolant 58 through the cooling passages 56 and out of the coolant outlets 64 at the tip 50 is primarily driven by a pressure difference defined between the supply pressure at the coolant passage inlets 60 and a static pressure which is typically defined at the tip 50 of the airfoil 40 , particularly within the tip pockets 74 , 76 . If the supply pressure is too low, for example, due to aerodynamic loading optimization or a decrease in operating speed or change in turbine load requirements, a lower static pressure is required in order to meet cooling flow needs.
  • a slot or opening 82 is formed along the suction side portion 70 of the tip rail 66 .
  • the slot 82 is formed along the suction side portion 70 of the tip rail 66 to define a coolant flow path 84 that provides for fluid communication of the coolant 58 from the first tip pocket 74 into the low or reduced pressure region 80 , thus reducing the static pressure within the first tip pocket 74 , thereby promoting or enhancing coolant flow through the airfoil 40 , particularly proximate to the leading edge 52 .
  • FIGS. 5 , 6 , 7 and 8 are enlarged perspective views of the tip 50 of the airfoil 40 , according to various embodiments of the present invention.
  • a slot or opening 86 may be defined along the suction side portion 70 of the tip rail 66 so as to define a coolant flow path 88 that provides for fluid communication of the coolant 58 from the second tip pocket 76 into the low or reduced pressure region 80 , thus reducing the static pressure with the second tip pocket 76 , thereby promoting or enhancing coolant flow through the airfoil 40 , particularly proximate to a mid-portion and/or the trialing edge 54 of the airfoil 40 .
  • the tip 50 may include slot 82 which defines a first slot and slot 86 which defines a second slot 88 where both slots 82 , 86 are defined along the suction side portion 70 of the tip rail 66 so as to define coolant flow paths 84 , 88 that provide for fluid communication of the coolant 58 from the first and second tip pockets 74 , 76 respectively into the low or reduced pressure region 80 , thus reducing the static pressure within both the first and second tip pockets 74 , 76 , thereby promoting or enhancing coolant flow through the airfoil 40 , proximate to both the leading edge 52 , a mid-portion and/or the trialing edge 54 of the airfoil 40 .
  • the tip 50 may include a secondary baffle 90 which extends radially outwardly from the tip floor 62 and from the pressure side to the suction side portions 68 , 70 of the tip rail 66 , thus defining a third tip pocket 92 .
  • a slot 94 may be defined along the suction side portion 70 of the tip rail 66 so as to define a coolant flow path 96 that provides for fluid communication of the coolant 58 from the third tip pocket 92 into the low or reduced pressure region 80 , thus reducing the static pressure within the third tip pocket 92 , thereby promoting or enhancing coolant flow through the airfoil 40 , proximate to the trialing edge 54 of the airfoil 40 .
  • the baffle 72 is configured to allow at least a portion of the coolant 58 to flow between the first and second tip pockets 74 , 76 .
  • a slot or opening 96 may be formed along the baffle 72 , thus defining a flow path 98 therebetween.
  • the baffle 72 may be sized from the tip floor 62 such that at least a portion of the coolant 58 that flows into the first tip pocket 74 is allowed to flow over a top portion 98 of the baffle 72 , thus reducing the static pressure in the first tip pocket 74 , thereby promoting or enhancing coolant flow through the airfoil 40 , proximate to the leading edge 52 of the airfoil 40 .
  • the present invention provides various technical benefits over existing rotor blade tip technologies.
  • the present invention provides lower static pressures for various cooling flows, especially for flows along the leading edge of the rotor blade airfoil.
  • the lower static pressures are achieved by segregating the tip into separate tip pockets or regions, and connecting the different tip pockets with different pressure zones.
  • the reduced static pressure at the tip may reduce the required coolant supply pressure at the cooling passage inlets, thus resulting in improved overall turbine performance.

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  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
US14/289,908 2014-05-29 2014-05-29 Rotor blade cooling flow Abandoned US20150345301A1 (en)

Priority Applications (5)

Application Number Priority Date Filing Date Title
US14/289,908 US20150345301A1 (en) 2014-05-29 2014-05-29 Rotor blade cooling flow
DE102015107844.9A DE102015107844A1 (de) 2014-05-29 2015-05-19 Rotorblattkühlströmung
JP2015103298A JP2015224634A (ja) 2014-05-29 2015-05-21 ロータブレードクーラント流
CH00756/15A CH709650A2 (de) 2014-05-29 2015-05-27 Rotorblatt mit Kühldurchgängen.
CN201510285010.9A CN105317468A (zh) 2014-05-29 2015-05-29 转子叶片冷却流

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
US14/289,908 US20150345301A1 (en) 2014-05-29 2014-05-29 Rotor blade cooling flow

Publications (1)

Publication Number Publication Date
US20150345301A1 true US20150345301A1 (en) 2015-12-03

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ID=54481598

Family Applications (1)

Application Number Title Priority Date Filing Date
US14/289,908 Abandoned US20150345301A1 (en) 2014-05-29 2014-05-29 Rotor blade cooling flow

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US (1) US20150345301A1 (ja)
JP (1) JP2015224634A (ja)
CN (1) CN105317468A (ja)
CH (1) CH709650A2 (ja)
DE (1) DE102015107844A1 (ja)

Cited By (7)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US20170145827A1 (en) * 2015-11-23 2017-05-25 United Technologies Corporation Turbine blade with airfoil tip vortex control
US20180202297A1 (en) * 2012-07-03 2018-07-19 United Technologies Corporation Tip leakage flow directionality control
US20180258774A1 (en) * 2012-07-03 2018-09-13 United Technologies Corporation Tip leakage flow directionality control
CN108868894A (zh) * 2017-05-10 2018-11-23 通用电气公司 转子叶片及对应的燃气涡轮
US20190003317A1 (en) * 2017-06-30 2019-01-03 General Electric Company Turbomachine rotor blade
US10677066B2 (en) 2015-11-23 2020-06-09 United Technologies Corporation Turbine blade with airfoil tip vortex control
EP3597859B1 (en) * 2018-07-13 2023-08-30 Honeywell International Inc. Turbine blade with dust tolerant cooling system

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US11434770B2 (en) 2017-03-28 2022-09-06 Raytheon Technologies Corporation Tip cooling design
US20180320530A1 (en) * 2017-05-05 2018-11-08 General Electric Company Airfoil with tip rail cooling
KR102155797B1 (ko) * 2019-04-15 2020-09-14 두산중공업 주식회사 터빈 블레이드 및 이를 포함하는 터빈

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US4761116A (en) * 1987-05-11 1988-08-02 General Electric Company Turbine blade with tip vent
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US20020162220A1 (en) * 2001-05-07 2002-11-07 Kevin Updegrove Method of repairing a turbine blade tip
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US8083484B2 (en) * 2008-12-26 2011-12-27 General Electric Company Turbine rotor blade tips that discourage cross-flow
US20120201695A1 (en) * 2009-06-17 2012-08-09 Little David A Turbine blade squealer tip rail with fence members
US20130039773A1 (en) * 2011-08-12 2013-02-14 Stanley J. Funk Method of measuring turbine blade tip erosion
US8435004B1 (en) * 2010-04-13 2013-05-07 Florida Turbine Technologies, Inc. Turbine blade with tip rail cooling
US20150330228A1 (en) * 2014-05-16 2015-11-19 United Technologies Corporation Airfoil tip pocket with augmentation features

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EP2725195B1 (en) * 2012-10-26 2019-09-25 Rolls-Royce plc Turbine blade and corresponding rotor stage

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US4589824A (en) * 1977-10-21 1986-05-20 United Technologies Corporation Rotor blade having a tip cap end closure
US4606701A (en) * 1981-09-02 1986-08-19 Westinghouse Electric Corp. Tip structure for a cooled turbine rotor blade
US4761116A (en) * 1987-05-11 1988-08-02 General Electric Company Turbine blade with tip vent
US5738491A (en) * 1997-01-03 1998-04-14 General Electric Company Conduction blade tip
US20020162220A1 (en) * 2001-05-07 2002-11-07 Kevin Updegrove Method of repairing a turbine blade tip
US20090180895A1 (en) * 2008-01-10 2009-07-16 General Electric Company Turbine blade tip shroud
US8083484B2 (en) * 2008-12-26 2011-12-27 General Electric Company Turbine rotor blade tips that discourage cross-flow
US20120201695A1 (en) * 2009-06-17 2012-08-09 Little David A Turbine blade squealer tip rail with fence members
US8435004B1 (en) * 2010-04-13 2013-05-07 Florida Turbine Technologies, Inc. Turbine blade with tip rail cooling
US20130039773A1 (en) * 2011-08-12 2013-02-14 Stanley J. Funk Method of measuring turbine blade tip erosion
US20150330228A1 (en) * 2014-05-16 2015-11-19 United Technologies Corporation Airfoil tip pocket with augmentation features

Cited By (10)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US20180202297A1 (en) * 2012-07-03 2018-07-19 United Technologies Corporation Tip leakage flow directionality control
US20180258774A1 (en) * 2012-07-03 2018-09-13 United Technologies Corporation Tip leakage flow directionality control
US10774659B2 (en) * 2012-07-03 2020-09-15 Raytheon Technologies Corporation Tip leakage flow directionality control
US10815790B2 (en) * 2012-07-03 2020-10-27 Raytheon Technologies Corporation Tip leakage flow directionality control
US20170145827A1 (en) * 2015-11-23 2017-05-25 United Technologies Corporation Turbine blade with airfoil tip vortex control
US10677066B2 (en) 2015-11-23 2020-06-09 United Technologies Corporation Turbine blade with airfoil tip vortex control
CN108868894A (zh) * 2017-05-10 2018-11-23 通用电气公司 转子叶片及对应的燃气涡轮
US20190003317A1 (en) * 2017-06-30 2019-01-03 General Electric Company Turbomachine rotor blade
US10590777B2 (en) * 2017-06-30 2020-03-17 General Electric Company Turbomachine rotor blade
EP3597859B1 (en) * 2018-07-13 2023-08-30 Honeywell International Inc. Turbine blade with dust tolerant cooling system

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DE102015107844A1 (de) 2015-12-03
JP2015224634A (ja) 2015-12-14
CH709650A2 (de) 2015-11-30
CN105317468A (zh) 2016-02-10

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