US20120045337A1 - Turbine bucket assembly and methods for assembling same - Google Patents

Turbine bucket assembly and methods for assembling same Download PDF

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Publication number
US20120045337A1
US20120045337A1 US12/860,493 US86049310A US2012045337A1 US 20120045337 A1 US20120045337 A1 US 20120045337A1 US 86049310 A US86049310 A US 86049310A US 2012045337 A1 US2012045337 A1 US 2012045337A1
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US
United States
Prior art keywords
sealing
sealing assembly
cover plate
rotor blade
rotor
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
US12/860,493
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English (en)
Inventor
Michael James Fedor
David Martin Johnson
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 US12/860,493 priority Critical patent/US20120045337A1/en
Assigned to GENERAL ELECTRIC COMPANY reassignment GENERAL ELECTRIC COMPANY ASSIGNMENT OF ASSIGNORS INTEREST (SEE DOCUMENT FOR DETAILS). Assignors: FEDOR, MICHAEL JAMES, JOHNSON, DAVID MARTIN
Priority to DE102011052419A priority patent/DE102011052419A1/de
Priority to CH01333/11A priority patent/CH703658A2/de
Priority to JP2011178178A priority patent/JP2012041930A/ja
Priority to CN2011102802130A priority patent/CN102373961A/zh
Publication of US20120045337A1 publication Critical patent/US20120045337A1/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
    • F01D11/00—Preventing or minimising internal leakage of working-fluid, e.g. between stages
    • F01D11/005—Sealing means between non relatively rotating elements
    • F01D11/006—Sealing the gap between rotor blades or blades and rotor
    • 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
    • Y10—TECHNICAL SUBJECTS COVERED BY FORMER USPC
    • Y10T—TECHNICAL SUBJECTS COVERED BY FORMER US CLASSIFICATION
    • Y10T29/00—Metal working
    • Y10T29/49—Method of mechanical manufacture
    • Y10T29/49316—Impeller making
    • Y10T29/4932—Turbomachine making
    • Y10T29/49321—Assembling individual fluid flow interacting members, e.g., blades, vanes, buckets, on rotary support member

Definitions

  • the subject matter described herein relates generally to gas turbine engines and, more particularly, to a bucket assembly for use with a turbine engine.
  • At least some known rotor assemblies used with turbine engines include at least one row of circumferentially-spaced rotor blades.
  • Each rotor blade includes an airfoil that includes a pressure side and a suction side that are connected together along leading and trailing edges.
  • Each airfoil extends radially outward from a rotor blade platform.
  • Each rotor blade also includes a dovetail that extends radially inward from a shank defined between the platform and the dovetail. The dovetail is used to mount the rotor blade to a rotor disk or spool.
  • Known blades are hollow and include an internal cooling cavity that is defined at least partially by the airfoil, platform, shank, and dovetail and that is used to channel a flow of cooling fluid. Leakage of cooling fluid may occur between adjacent rotor blades. Depending on the amount of leakage, turbine performance and output may be adversely impacted.
  • the airfoil portions of at least some known rotor blades are generally exposed to higher temperatures than the dovetail portions. Higher temperatures may cause temperature mismatches to develop at the interface between the airfoil and the platform, and/or between the shank and the platform. These temperature mismatches may cause compressive thermal stresses to be induced to the rotor blade platform. Over time, continued operation with high compressive thermal stresses may cause platform oxidation, platform cracking, and/or platform creep deflection, any or all of which may shorten the useful life of the rotor assembly.
  • a method for assembling a rotor assembly for use with a turbine engine includes providing at least two rotor blades that each include a shank extending between a dovetail and a platform.
  • the shank includes at least one cover plate that extends inwardly from the platform towards the dovetail.
  • An airfoil extends outwardly from the platform.
  • a first rotor blade is coupled to a rotor disk.
  • a second rotor blade is coupled to the rotor disk, such that a cavity is defined between the first and second rotor blades, and such that a seal path is defined between a first rotor blade cover plate and a second rotor blade cover plate.
  • a rotor blade for a turbine engine includes a platform that includes a radially outer surface and a radially inner surface.
  • An airfoil extends radially outwardly from the platform.
  • a dovetail is adapted to be coupled to a rotor wheel.
  • a shank extends between the platform and the dovetail.
  • the shank includes at least one cover plate that extends inwardly from the platform towards the dovetail.
  • At least one sealing assembly is coupled to the cover plate. The sealing assembly extends from the dovetail to the platform. The sealing assembly forms a seal path between the rotor blade and a circumferentially adjacent rotor blade.
  • a gas turbine engine in another aspect, includes a compressor and a combustor coupled downstream from the compressor to receive at least some of the air discharged by the compressor.
  • a rotor shaft is coupled to the compressor.
  • a plurality of circumferentially-spaced rotor blades are coupled to the rotor shaft.
  • Each of the plurality of rotor blades includes a platform.
  • An airfoil extends radially outwardly from the platform.
  • a dovetail is coupled to the rotor shaft.
  • a shank extends between the platform and the dovetail.
  • the shank includes at least one cover plate that extends inwardly from the platform towards the dovetail.
  • At least one sealing assembly is coupled to the cover plate such that a seal path is defined between adjacent rotor blades.
  • FIG. 1 is schematic illustration of an exemplary known turbine engine system.
  • FIG. 2 is an enlarged perspective view of an exemplary rotor assembly that may be used with the turbine engine system shown in FIG. 1 .
  • FIG. 3 is an enlarged sectional view of a portion of the rotor assembly shown in FIG. 2
  • FIG. 4 is a cross-sectional view of the rotor assembly shown in FIG. 2 .
  • the exemplary methods and systems described herein overcome disadvantages of known rotor blade assemblies by providing a rotor blade that facilitates reducing leakage of cooling fluid from the rotor blade. More specifically, the embodiments described herein include a labyrinth seal path that is positioned between adjoining rotor blades to facilitate increasing a back pressure between adjacent rotor blades and to facilitate reducing leakage of cooling fluid through the rotor blades.
  • rotor blade is used interchangeably with the term “bucket” and thus can include any combination of a bucket including a platform and dovetail and/or a bucket integrally formed with the rotor disk, either of which may include at least one airfoil segment.
  • FIG. 1 is a schematic view of an exemplary gas turbine engine 10 .
  • gas turbine engine 10 includes an intake section 12 , a compressor section 14 coupled downstream from intake section 12 , a combustor section 16 coupled downstream from compressor section 14 , a turbine section 18 coupled downstream from combustor section 16 , and an exhaust section 20 .
  • Turbine section 18 is includes a rotor assembly 22 that is coupled to compressor section 14 via a drive shaft 32 .
  • Combustor section 16 includes a plurality of combustors 24 .
  • Combustor section 16 is coupled to compressor section 14 such that each combustor 24 is in flow communication with compressor section 14 and such that fuel nozzle assembly 26 is coupled to each combustor 24 .
  • Turbine section 18 is rotatably coupled to compressor section 14 and to a load 28 such as, but not limited to, an electrical generator and a mechanical drive application.
  • compressor section 14 and turbine section 18 each include at least one turbine blade or bucket 30 coupled to rotor assembly 22 that include airfoil portions (not shown in FIG. 1 ).
  • intake section 12 channels air towards compressor section 14 .
  • Compressor section 14 compresses the inlet air to a higher pressure and temperature and discharges the compressed air towards combustor section 16 .
  • the compressed air is mixed with fuel and ignited to generate combustion gases that flow to turbine section 18 .
  • Turbine section 18 drives compressor section 14 and/or load 28 .
  • Fuel is channeled to fuel nozzle assembly 26 wherein it is mixed with the air and ignited in combustor section 16 .
  • Combustion gases are generated and channeled to turbine section 18 wherein gas stream thermal energy is converted to mechanical rotational energy. Exhaust gases exit turbine section 18 and flow through exhaust section 20 to ambient atmosphere.
  • FIG. 2 is an enlarged perspective view of an exemplary rotor assembly 22 that may be used with gas turbine engine 10 (shown in FIG. 1 ).
  • FIG. 3 is an enlarged sectional view of a portion of rotor assembly 22
  • FIG. 4 is a cross-sectional view of rotor assembly 22 taken along sectional line 4 - 4 in FIG. 3 .
  • rotor assembly 22 includes at least one rotor blade 100 coupled to a rotor disk 102 .
  • rotor assembly 22 includes a first rotor blade 104 , a second rotor blade 106 , and at least a third rotor blade 107 .
  • each rotor blade 100 is coupled to a rotor disk 102 that is rotatably coupled to a rotor shaft, such as drive shaft 32 (shown in FIG. 1 ).
  • rotor blades 100 are mounted within a rotor spool (not shown). More specifically, when rotor blades 100 are coupled to rotor disk 102 , a gap 108 is defined between adjacent circumferentially-spaced rotor blades 100 .
  • each rotor blade 100 extends radially outward from rotor disk 102 and includes an airfoil 110 , a platform 112 , a shank 114 , and a dovetail 116 .
  • Each airfoil 110 includes a first sidewall 118 and a second sidewall 120 that is coupled to first sidewall 118 to form airfoil 110 .
  • first sidewall 118 is convex and defines a suction side 119 of airfoil 110
  • second sidewall 120 is concave and defines a pressure side 121 of airfoil 110
  • First sidewall 118 is coupled to second sidewall 120 along a leading edge 122 and along an axially-spaced trailing edge 124 of airfoil 110 . More specifically, airfoil trailing edge 124 is spaced chord-wise and downstream from airfoil leading edge 122 .
  • First sidewall 118 and second sidewall 120 each extend longitudinally or radially outwardly in span from a blade root 126 positioned adjacent to platform 112 , to an airfoil tip 128 .
  • an internal cooling chamber 130 is defined within airfoil 110 between first sidewall 118 and second sidewall 120 , and extends through platform 112 , through shank 114 , and into dovetail 116 .
  • Platform 112 extends between airfoil 110 and shank 114 such that each airfoil 110 extends radially outwardly from platform 112 .
  • Shank 114 extends radially inwardly from platform 112 to dovetail 116 .
  • Dovetail 116 extends radially inwardly from shank 114 to enable rotor blades 100 to be coupled to rotor disk 102 .
  • Platform 112 includes an upstream side or skirt 132 , and a downstream side or skirt 134 that are connected together with a pressure-side edge 136 and an opposite suction-side edge 138 .
  • a gap 140 is defined between circumferentially adjacent rotor blade platforms 112 , and more specifically between pressure-side edge 136 and an adjacent suction-side edge 138 .
  • shank 114 includes a first sidewall 142 , a second sidewall 144 , an upstream sidewall or forward cover plate 146 , and an opposite downstream sidewall or aft cover plate 148 .
  • first sidewall 142 is substantially concave and is coupled between forward cover plate 146 and aft cover plate 148 such that forward cover plate 146 is opposite aft cover plate 148 .
  • Second sidewall 144 is substantially convex and is coupled between forward cover plate 146 and aft cover plate 148 .
  • first sidewall 142 is coupled to second sidewall 144 such that a cavity 150 is defined at least partially between first sidewall 142 and second sidewall 144 .
  • first sidewall 142 is coupled to second sidewall 144 such that a unitary member extending between forward cover plate 146 and aft cover plate 148 is formed.
  • shank 114 is formed as a unitary member.
  • first sidewall 142 and second sidewall 144 are each recessed with respect to forward cover plate 146 and aft cover plate 148 , respectively, such that when rotor blades 100 are coupled to rotor disk 102 , a shank cavity 152 is defined between first sidewall 142 and an adjacent second sidewall 144 .
  • a forward angel wing 154 extends outwardly from forward cover plate 146 .
  • An aft angel wing 156 extends outwardly from aft cover plate 148 .
  • Forward angel wing 154 and aft angel wing 156 each facilitate sealing forward and aft angel wing buffer cavities (not shown) defined within rotor assembly 22 .
  • a forward lower angel wing 158 extends outwardly from forward cover plate 146 , and is configured to facilitate sealing between rotor blade 100 and rotor disk 102 . More specifically, forward lower angel wing 158 extends outwardly from forward cover plate 146 between dovetail 116 and forward angel wing 154 .
  • aft cover plate 148 includes a leading edge portion 164 and a circumferentially-spaced trailing edge portion 166 .
  • a first sealing assembly 168 is coupled to leading edge portion 164
  • a second sealing assembly 170 is coupled to trailing edge portion 166 .
  • first sealing assembly 168 cooperates with an adjacent second sealing assembly 170 when rotor blades 100 are coupled to rotor disk 102 .
  • First sealing assembly 168 and second sealing assembly 170 each extend between dovetail 116 and platform 112 , and each facilitates sealing shank cavity 152 .
  • first sealing assembly 168 and second sealing assembly 170 cooperate to form a seal path 172 between a first aft cover plate 148 and an adjacent second aft cover plate 148 .
  • Seal path 172 facilitates reducing a volume of air channeled between circumferentially adjacent rotor blade shanks 114 . More specifically, seal path 172 facilitates reducing the volume of air that must be channeled from forward cover plate 146 to aft cover plate 148 through shank cavity 152 to facilitate preventing a flow of hot gases from entering shank cavity 152 .
  • aft cover plate 148 extends a radial height r 1 from dovetail 116 to a platform inner surface 174 .
  • First sealing assembly 168 and second sealing assembly 170 each extend a radial height r 2 from dovetail 116 to platform inner surface 174 .
  • Radial height r 2 is approximately the same height as radial height r 1 of aft cover plate 148 .
  • first sealing assembly 168 and/or second sealing assembly 170 extends the full radial height r 1 of aft cover plate 148 .
  • first sealing assembly 168 includes a sealing extension 176 that extends outwardly from leading edge portion 164 towards an adjacent rotor blade trailing edge portion 166 .
  • Second sealing assembly 170 includes a recessed sealing groove 178 that is defined within trailing edge portion 166 . Recessed sealing groove 178 is sized to receive an adjacent sealing extension 176 such that recessed sealing groove 178 and sealing extension 176 cooperate to form seal path 172 .
  • first sealing assembly 168 includes recessed sealing groove 178 and second sealing assembly 170 includes sealing extension 176 .
  • first rotor blade 104 includes first sealing assembly 168 , including sealing extension 176 , and second sealing assembly 170 , including recessed sealing groove 178 .
  • first rotor blade 104 includes first sealing assembly 168 , including recessed sealing groove 178 , and second sealing assembly 170 , including a sealing extension 176 .
  • second rotor blade 106 includes first sealing assembly 168 and second sealing assembly 170 each including sealing extension 176 .
  • second rotor blade 106 includes first sealing assembly 168 and second sealing assembly 170 each including recessed sealing groove 178 .
  • recessed sealing groove 178 includes a radially outer surface 184 that extends between dovetail 116 and platform inner surface 174 .
  • An abradable layer 186 is coupled to recessed sealing groove outer surface 184 .
  • abradable layer 186 includes an aluminum composite material.
  • sealing extension 176 includes a plurality of labyrinth teeth 188 that extend outwardly from an inner surface 190 of sealing extension 176 .
  • Labyrinth teeth 188 are each positioned adjacent to an opposing recessed sealing groove outer surface 184 such that a labyrinth seal 191 is defined between sealing extension 176 and recessed sealing groove 178 .
  • shank 114 includes a leading edge radial seal pin slot 192 that extends generally radially through shank 114 at least partially between platform 112 and dovetail 116 . More specifically, leading edge radial seal pin slot 192 is defined within shank forward cover plate 146 and is adjacent to shank convex sidewall 144 . Leading edge radial seal pin slot 192 is sized to receive a radial seal pin 194 to facilitate sealing between adjacent forward cover plates 146 when rotor blades 100 are coupled within rotor disk 102 . In one embodiment, radial seal pin 194 is not inserted into leading edge radial seal pin slot 192 . In an alternative embodiment, forward cover plate 146 includes a first sealing assembly 168 and a second sealing assembly 170 .
  • combustion gases 196 contact rotor blades 100 causing rotor assembly 22 to rotate about drive shaft 32 .
  • At least a portion of combustion gases 196 pass through adjacent forward cover plates 146 , around radial seal pin 194 , and into shank cavity 152 .
  • First sealing assembly 168 and second sealing assembly 170 each facilitate preventing combustion gases 196 from passing through adjacent aft cover plates 148 causing an increase in a fluid pressure within shank cavity 152 that facilitates reducing a volume of combustion gases 196 entering shank cavity 152 .
  • the above-described methods and apparatus facilitate reducing an operating temperature of a rotor assembly. More specifically, the labyrinth seal defined between adjacent rotor blades facilitate reducing leakage of cooling fluid between adjacent rotor blades. In addition, the embodiments described herein facilitate increasing a back pressure of cooling fluid within a shank cavity, which facilitates increasing a flow of cooling fluid to the rotor blades to reduce an operating temperature of the rotor assembly. As such, the cost of maintaining the gas turbine engine system is facilitated to be reduced.
  • Exemplary embodiments of methods and apparatus for a turbine bucket assembly are described above in detail.
  • the methods and apparatus are not limited to the specific embodiments described herein, but rather, components of systems and/or steps of the method may be utilized independently and separately from other components and/or steps described herein.
  • the methods and apparatus may also be used in combination with other combustion systems and methods, and are not limited to practice with only the gas turbine engine assembly as described herein. Rather, the exemplary embodiment can be implemented and utilized in connection with many other combustion system applications.

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  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
  • Sealing Using Fluids, Sealing Without Contact, And Removal Of Oil (AREA)
US12/860,493 2010-08-20 2010-08-20 Turbine bucket assembly and methods for assembling same Abandoned US20120045337A1 (en)

Priority Applications (5)

Application Number Priority Date Filing Date Title
US12/860,493 US20120045337A1 (en) 2010-08-20 2010-08-20 Turbine bucket assembly and methods for assembling same
DE102011052419A DE102011052419A1 (de) 2010-08-20 2011-08-04 Turbinenschaufelanordnung und Verfahren zur Montage derselben
CH01333/11A CH703658A2 (de) 2010-08-20 2011-08-11 Laufschaufel mit einer Dichtungsanordnung zu benachbarter Laufschaufel.
JP2011178178A JP2012041930A (ja) 2010-08-20 2011-08-17 タービンバケット組立体及びそれを組立てる方法
CN2011102802130A CN102373961A (zh) 2010-08-20 2011-08-19 涡轮动叶组件和用于组装其的方法

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
US12/860,493 US20120045337A1 (en) 2010-08-20 2010-08-20 Turbine bucket assembly and methods for assembling same

Publications (1)

Publication Number Publication Date
US20120045337A1 true US20120045337A1 (en) 2012-02-23

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US12/860,493 Abandoned US20120045337A1 (en) 2010-08-20 2010-08-20 Turbine bucket assembly and methods for assembling same

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US (1) US20120045337A1 (de)
JP (1) JP2012041930A (de)
CN (1) CN102373961A (de)
CH (1) CH703658A2 (de)
DE (1) DE102011052419A1 (de)

Cited By (6)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US20130052020A1 (en) * 2011-08-23 2013-02-28 General Electric Company Coupled blade platforms and methods of sealing
CN103362561A (zh) * 2012-03-29 2013-10-23 通用电气公司 流路附近密封件隔离鸠尾榫
WO2014008015A1 (en) * 2012-07-02 2014-01-09 United Technologies Corporation Gas turbine engine component having platform cooling channel
EP3498980A1 (de) * 2017-12-15 2019-06-19 Ansaldo Energia Switzerland AG Stufenfalzdichtungsanordnung
US10648354B2 (en) * 2016-12-02 2020-05-12 Honeywell International Inc. Turbine wheels, turbine engines including the same, and methods of forming turbine wheels with improved seal plate sealing
US10851661B2 (en) 2017-08-01 2020-12-01 General Electric Company Sealing system for a rotary machine and method of assembling same

Families Citing this family (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
JP4999277B2 (ja) * 2005-02-10 2012-08-15 株式会社リコー 光学素子の固定接合方法、固定接合装置、光走査装置及び画像形成装置
DE102013220467A1 (de) * 2013-10-10 2015-05-07 MTU Aero Engines AG Rotor mit einem Rotorgrundkörper und einer Mehrzahl daran angebrachter Laufschaufeln
KR102525225B1 (ko) * 2021-03-12 2023-04-24 두산에너빌리티 주식회사 터보머신

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US6190131B1 (en) * 1999-08-31 2001-02-20 General Electric Co. Non-integral balanced coverplate and coverplate centering slot for a turbine
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US7175387B2 (en) * 2001-09-25 2007-02-13 Alstom Technology Ltd. Seal arrangement for reducing the seal gaps within a rotary flow machine
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US20080181767A1 (en) * 2007-01-30 2008-07-31 Siemens Power Generation, Inc. Turbine seal plate locking system
US20080196247A1 (en) * 2007-02-15 2008-08-21 Srinivas Ravi Method and apparatus to facilitate increasing turbine rotor efficiency

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DE69503798T2 (de) * 1994-10-31 1999-01-14 Westinghouse Electric Corp Gasturbinenschaufel mit gekühlter schaufelplattform
JP3462695B2 (ja) * 1997-03-12 2003-11-05 三菱重工業株式会社 ガスタービン動翼シール板
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US7322796B2 (en) * 2005-08-31 2008-01-29 United Technologies Corporation Turbine vane construction
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US3137478A (en) * 1962-07-11 1964-06-16 Gen Electric Cover plate assembly for sealing spaces between turbine buckets
US3807898A (en) * 1970-03-14 1974-04-30 Secr Defence Bladed rotor assemblies
US3644058A (en) * 1970-05-18 1972-02-22 Westinghouse Electric Corp Axial positioner and seal for turbine blades
US4422827A (en) * 1982-02-18 1983-12-27 United Technologies Corporation Blade root seal
US4507052A (en) * 1983-03-31 1985-03-26 General Motors Corporation End seal for turbine blade bases
US4669959A (en) * 1984-07-23 1987-06-02 United Technologies Corporation Breach lock anti-rotation key
US6190131B1 (en) * 1999-08-31 2001-02-20 General Electric Co. Non-integral balanced coverplate and coverplate centering slot for a turbine
US7175387B2 (en) * 2001-09-25 2007-02-13 Alstom Technology Ltd. Seal arrangement for reducing the seal gaps within a rotary flow machine
US20060073021A1 (en) * 2004-10-06 2006-04-06 Siemens Westinghouse Power Corporation Remotely accessible locking system for turbine blades
US20070258816A1 (en) * 2005-09-26 2007-11-08 Pratt & Whitney Canada Corp. Blades for a gas turbine engine with integrated sealing plate and method
US20080181767A1 (en) * 2007-01-30 2008-07-31 Siemens Power Generation, Inc. Turbine seal plate locking system
US20080196247A1 (en) * 2007-02-15 2008-08-21 Srinivas Ravi Method and apparatus to facilitate increasing turbine rotor efficiency

Cited By (13)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US20130052020A1 (en) * 2011-08-23 2013-02-28 General Electric Company Coupled blade platforms and methods of sealing
US8888459B2 (en) * 2011-08-23 2014-11-18 General Electric Company Coupled blade platforms and methods of sealing
US9151169B2 (en) 2012-03-29 2015-10-06 General Electric Company Near-flow-path seal isolation dovetail
CN103362561A (zh) * 2012-03-29 2013-10-23 通用电气公司 流路附近密封件隔离鸠尾榫
US9845687B2 (en) 2012-07-02 2017-12-19 United Technologies Corporation Gas turbine engine component having platform cooling channel
US9303518B2 (en) 2012-07-02 2016-04-05 United Technologies Corporation Gas turbine engine component having platform cooling channel
WO2014008015A1 (en) * 2012-07-02 2014-01-09 United Technologies Corporation Gas turbine engine component having platform cooling channel
US10053991B2 (en) 2012-07-02 2018-08-21 United Technologies Corporation Gas turbine engine component having platform cooling channel
US10648354B2 (en) * 2016-12-02 2020-05-12 Honeywell International Inc. Turbine wheels, turbine engines including the same, and methods of forming turbine wheels with improved seal plate sealing
US10851660B2 (en) * 2016-12-02 2020-12-01 Honeywell International Inc. Turbine wheels, turbine engines including the same, and methods of forming turbine wheels with improved seal plate sealing
US11015472B2 (en) 2016-12-02 2021-05-25 Honeywell International Inc. Turbine wheels, turbine engines including the same, and methods of forming turbine wheels with improved seal plate sealing
US10851661B2 (en) 2017-08-01 2020-12-01 General Electric Company Sealing system for a rotary machine and method of assembling same
EP3498980A1 (de) * 2017-12-15 2019-06-19 Ansaldo Energia Switzerland AG Stufenfalzdichtungsanordnung

Also Published As

Publication number Publication date
DE102011052419A1 (de) 2012-02-23
CN102373961A (zh) 2012-03-14
CH703658A2 (de) 2012-02-29
JP2012041930A (ja) 2012-03-01

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Owner name: GENERAL ELECTRIC COMPANY, NEW YORK

Free format text: ASSIGNMENT OF ASSIGNORS INTEREST;ASSIGNORS:FEDOR, MICHAEL JAMES;JOHNSON, DAVID MARTIN;REEL/FRAME:024866/0930

Effective date: 20100819

STCB Information on status: application discontinuation

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