US20130202439A1 - Rotating assembly for a turbine assembly - Google Patents

Rotating assembly for a turbine assembly Download PDF

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Publication number
US20130202439A1
US20130202439A1 US13/368,931 US201213368931A US2013202439A1 US 20130202439 A1 US20130202439 A1 US 20130202439A1 US 201213368931 A US201213368931 A US 201213368931A US 2013202439 A1 US2013202439 A1 US 2013202439A1
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United States
Prior art keywords
edge
substrate
rotating assembly
tip
hard
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
US13/368,931
Inventor
Kevin Leon Bruce
Ronald Ralph Cairo
Matthew Robert Piersall
Richard Gordon Rollings
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 US13/368,931 priority Critical patent/US20130202439A1/en
Assigned to GENERAL ELECTRIC COMPANY reassignment GENERAL ELECTRIC COMPANY ASSIGNMENT OF ASSIGNORS INTEREST (SEE DOCUMENT FOR DETAILS). Assignors: ROLLINGS, RICHARD GORDON, BRUCE, KEVIN LEON, Piersall, Matthew Robert, CAIRO, RONALD RALPH
Publication of US20130202439A1 publication Critical patent/US20130202439A1/en
Abandoned legal-status Critical Current

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    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01DNON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D11/00Preventing or minimising internal leakage of working-fluid, e.g. between stages
    • F01D11/08Preventing or minimising internal leakage of working-fluid, e.g. between stages for sealing space between rotor blade tips and stator
    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01DNON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D11/00Preventing or minimising internal leakage of working-fluid, e.g. between stages
    • F01D11/08Preventing or minimising internal leakage of working-fluid, e.g. between stages for sealing space between rotor blade tips and stator
    • F01D11/12Preventing or minimising internal leakage of working-fluid, e.g. between stages for sealing space between rotor blade tips and stator using a rubstrip, e.g. erodible. deformable or resiliently-biased part
    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01DNON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D11/00Preventing or minimising internal leakage of working-fluid, e.g. between stages
    • F01D11/08Preventing or minimising internal leakage of working-fluid, e.g. between stages for sealing space between rotor blade tips and stator
    • F01D11/12Preventing or minimising internal leakage of working-fluid, e.g. between stages for sealing space between rotor blade tips and stator using a rubstrip, e.g. erodible. deformable or resiliently-biased part
    • F01D11/122Preventing or minimising internal leakage of working-fluid, e.g. between stages for sealing space between rotor blade tips and stator using a rubstrip, e.g. erodible. deformable or resiliently-biased part with erodable or abradable material
    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01DNON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D5/00Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/12Blades
    • F01D5/22Blade-to-blade connections, e.g. for damping vibrations
    • F01D5/225Blade-to-blade connections, e.g. for damping vibrations by shrouding
    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05DINDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2230/00Manufacture
    • F05D2230/20Manufacture essentially without removing material
    • F05D2230/22Manufacture essentially without removing material by sintering
    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05DINDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2250/00Geometry
    • F05D2250/20Three-dimensional
    • F05D2250/23Three-dimensional prismatic
    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05DINDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2250/00Geometry
    • F05D2250/20Three-dimensional
    • F05D2250/28Three-dimensional patterned
    • F05D2250/282Three-dimensional patterned cubic pattern
    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05DINDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2250/00Geometry
    • F05D2250/20Three-dimensional
    • F05D2250/28Three-dimensional patterned
    • F05D2250/283Three-dimensional patterned honeycomb
    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05DINDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2250/00Geometry
    • F05D2250/20Three-dimensional
    • F05D2250/29Three-dimensional machined; miscellaneous
    • F05D2250/294Three-dimensional machined; miscellaneous grooved
    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05DINDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2250/00Geometry
    • F05D2250/70Shape
    • F05D2250/75Shape given by its similarity to a letter, e.g. T-shaped

Abstract

A rotating assembly for a turbine assembly includes an airfoil extending radially outward from, and rotatable about, an axial centerline. Also included is a tip shroud integrally connected proximate a radially outer tip of the airfoil. Further included is a substrate operably coupled to the tip shroud. Yet further included is at least one hard face interface member secured to the substrate.

Description

    BACKGROUND OF THE INVENTION
  • The subject matter disclosed herein relates to turbine systems, and more particularly to tip shrouds.
  • Turbine systems employ a number of rotating components or assemblies, such as compressor stages and turbine stages that rotate at high speed when the turbine is in operation, for example. In general, a stage includes a plurality of free-floating blades that extend radially outward from a central hub. Some blades include a shroud that limits vibration within a stage. The shroud is typically positioned at a tip portion of the blade, a mid portion of the blade or at both the mid portion and the tip portion of the blade. The shrouds are designed such that at high or operational speeds, the free-floating blades interlock to form an integral rotating member, however, even at lower speeds and possibly even when starting from a starting position at 0 rpm, the blades may interlock. During blade interlock, wear occurs due to slip when the tip portions are interlocked. At lower speeds, such as on turbine turning gear, the blades may not interlock and will often times impact one another. Impacts between the blades can cause damage that will shorten service life of the turbomachine.
  • In order to minimize damage resulting from blade impacts, a hard face coating is applied to potential contact points. The hard face coating increases wear resistance that may occur during operation of the blades and also increases the durability of potential contact points that may be susceptible to impacts. Conventionally, the hard face coating is metallurgically bonded to the blade through, for example, a welding, brazing or spraying process. Using a welding process to bond the hard face interface to the blade inherently produces a great deal of localized heat which, if not properly controlled, can weaken the wear and impact resistance and other metallurgical properties at the interface of the materials being joined. Excessive heat can also cause cracking in adjacent material during manufacture.
  • BRIEF DESCRIPTION OF THE INVENTION
  • According to one aspect of the invention, a rotating assembly for a turbine assembly includes an airfoil extending radially outward from, and rotatable about, an axial centerline. Also included is a tip shroud integrally connected proximate a radially outer tip of the airfoil. Further included is a substrate operably coupled to the tip shroud. Yet further included is at least one hard face interface member secured to the substrate.
  • According to another aspect of the invention, a rotating assembly for a turbine assembly includes a plurality of rotating members extending radially outward from, and rotatable about, an axial centerline. Also included is a tip shroud integrally connected proximate a radially outer tip of at least one of the plurality of rotating members. Further included is a substrate operably coupled to the tip shroud. Yet further included is a hard face interface member having a base portion and a plurality of edge portions, wherein the base portion is secured to the substrate and a portion of the substrate at least partially surrounds at least one of the plurality of edge portions.
  • According to yet another aspect of the invention, a rotating assembly for a turbine assembly includes an airfoil extending radially outward from, and rotatable about, an axial centerline. Also included is a tip shroud integrally connected proximate a radially outer tip of the airfoil. Further included is a substrate operably coupled to the tip shroud, wherein the substrate includes a plurality of grooved portions. Yet further included is at least one hard face interface member being both mechanically interlocked with, and metallurgically bonded to the substrate.
  • These and other advantages and features will become more apparent from the following description taken in conjunction with the drawings.
  • BRIEF DESCRIPTION OF THE DRAWING
  • The subject matter, which is regarded as the invention, is particularly pointed out and distinctly claimed in the claims at the conclusion of the specification. The foregoing and other features and advantages of the invention are apparent from the following detailed description taken in conjunction with the accompanying drawings in which:
  • FIG. 1 is a partial, cross-sectional schematic view of a turbomachine including a rotating assembly;
  • FIG. 2 is a partial perspective view of a rotating assembly including a plurality of rotating components;
  • FIG. 3 is a front perspective view of a tip shroud of a first embodiment;
  • FIG. 4 is a rear, elevational view of the tip shroud of FIG. 3;
  • FIG. 5 is a perspective view of a substrate having a plurality of substantially “U” shaped grooves;
  • FIG. 6 is a perspective view of a substrate having a plurality of “V” shaped grooves; and
  • FIG. 7 is a front perspective view of a tip shroud of a second embodiment.
  • The detailed description explains embodiments of the invention, together with advantages and features, by way of example with reference to the drawings.
  • DETAILED DESCRIPTION OF THE INVENTION
  • Referring to FIG. 1, a turbomachine, shown in the form of a gas turbine engine, constructed in accordance with an exemplary embodiment of the present invention is indicated generally at 10. The engine 10 includes a compressor 12 and a plurality of combustor assemblies arranged in a can annular array, one of which is indicated at 14. As shown, the combustor assembly 14 includes an endcover assembly 16 that seals, and at least partially defines, a combustion chamber 18. A plurality of nozzles 20-22 are supported by the endcover assembly 16 and extend into the combustion chamber 18. The nozzles 20-22 receive fuel through a common fuel inlet (not shown) and compressed air from the compressor 12. The fuel and compressed air are passed into the combustion chamber 18 and ignited to form a high temperature, high pressure combustion product or air stream that is used to drive a turbine 24. The turbine 24 includes a plurality of rotating assemblies or stages 26-28 that are operationally connected to the compressor 12 through a compressor/turbine shaft 30 (sometimes referred to as a rotor).
  • In operation, air flows into the compressor 12 and is compressed into a high pressure gas. The high pressure gas is supplied to the combustor assembly 14 and mixed with fuel, for example process gas and/or synthetic gas (syngas), in the combustion chamber 18. The fuel/air or combustible mixture ignite to form a high pressure, high temperature combustion gas stream of approximately 538 degrees Celsius (° C.) to 1593° C. (1000 degrees Fahrenheit (° F.) to 2900° F.). Alternatively, the combustor assembly 14 can combust fuels that include, but are not limited to, natural gas and/or fuel oil. In any event, the combustor assembly 14 channels the combustion gas stream to the turbine 24 which converts thermal energy to mechanical, rotational energy.
  • At this point, it should be understood that each rotating assembly or stage 26-28 is similarly formed, thus reference will be made to FIG. 2 in describing stage 26 constructed in accordance with an exemplary embodiment of the present invention with an understanding that the remaining stages, i.e., stages 27 and 28, have corresponding structure. Also, it should be understood that the present invention could be employed in stages in the compressor 12 or other rotating assemblies that require wear and/or impact resistant surfaces. In any event, the stage 26 is shown to include a plurality of rotating members, such as an airfoil 32, which each extend radially outward from a central hub 34 having an axial centerline 35. The airfoil 32 is rotatable about the axial centerline 35 of the central hub 34 and includes a base portion 36 and a tip portion 38.
  • The tip shroud 50 covers a bucket or throat portion (not separately labeled) of airfoil 32. The tip shroud 50 is designed to receive, or nest with, tip shrouds on adjacent rotating members in order to form a continuous ring that extends circumferentially about the stage 26. The continuous ring creates an outer flow path boundary that reduces gas path air leakage over top portions (not separately labeled) of the stage 26, so as to increase stage efficiency and overall turbine performance. In the exemplary embodiment shown, during high or operational speeds, adjacent airfoils 32 interlock through their respective tip-shrouds 50 by virtue of centrifugal forces created by the operation of the turbine 24. It should be noted that interlock may occur even at extremely low speed operation, such that wearing due to slipping of the blades may occur due to operation during interlock. However, during lower speeds such as, during turbine turning gear, the rotational force may not be sufficient to establish the interlock and thus, adjacent rotating members may impact one another. The impacts can create wear on the rotating members thereby lowering an overall service life of the turbine 24. Additionally, operator handling at several manufacturing and assembly stages may result in such impacts. Towards that end, tip shroud 50 is provided with a wear resistant/impact resistant member in a manner that will be described more fully below.
  • Referring to FIGS. 3 and 4, the tip shroud 50 includes a tip blade 52 that extends generally along a top surface 54 of the tip shroud 50. The tip blade 52 is comprised of a first surface 56 and an opposing second surface 58, as well as a first edge 60 and a second edge 62. In the exemplary embodiment illustrated, the tip blade 52 also includes one or more cutter members, such as a first cutter member 64 and a second cutter member 66. The first cutter member 64 is disposed proximate the first edge 60 of the tip blade 52 and on the second surface 58, while the second cutter member 66 is disposed proximate the second edge 62 and on the first surface 56. Both the first and second cutter members 64, 66 are configured to engage a counterpart surface. An example of such a counterpart surface is an inner surface of an outer casing (not illustrated), such as a honeycomb structure. Engagement of the first and second cutter members 64, 66 with the counterpart surface removes material from the counterpart surface to achieve a sealing arrangement proximate the outer area of the stage 26.
  • Referring to FIGS. 5 and 6, a substrate 70 is operably coupled to the tip shroud 50 and includes a mechanical interlock pattern 72. In the illustrated embodiments, the mechanical interlock pattern 72 is shown as simply a plurality of grooves 74, but it is to be appreciated that the mechanical interlock pattern 72 may be formed of a plurality of grooves, protrusions, or a combination thereof. The plurality of grooves 74 may take on a number of geometries including, but not limited to, grooves having a relatively “U” shape (FIG. 5) or “V” shape (FIG. 6), for example. The “U” and “V” shaped examples are merely illustrative of the various shapes that may be employed to form the plurality of grooves 74. Irrespective of the mechanical interlock pattern 72, and specifically the groove or protrusion arrangement, the arrangement provides an increased surface area for a hard face interface 80 to interlock with upon application, when compared to a merely flat surface.
  • Referring again to FIGS. 3 and 4, the substrate 70 is shown as providing an interface with the tip shroud 50 and the hard face interface 80. The substrate 70 provides retention and/or bonding function for the hard face interface 80 to the tip shroud 50. The hard face interface 80 is operably secured to the substrate 70 by welding or brazing, for example. As previously described, the mechanical interlock pattern 72 of the substrate 70 increases the surface area available to form the securement of the hard face interface 80 to the substrate 70, thereby enhancing structural integrity. The hard face interface 80 includes a base portion 82 and a plurality of edge portions 84. It is contemplated that the substrate 70 may be secured to the hard face interface 80 at the base portion 82 or at least one of the plurality of edge portions 84, or a combination thereof. The substrate 70 provides additional structural integrity to the hard face interface 80 by partially or fully encapsulating one or more of the edge portions 84, such that the likelihood of damage to the hard face interface 80 due to a collision with other structures during operation, installation, or handling, is reduced. The hard face interface 80 may be formed of a pre-sintered preform (PSP) material, for example, but it is to be appreciated that various materials may be employed for differing applications of use.
  • Referring to FIG. 7, a tip shroud 50 similar to that illustrated in FIGS. 3 and 4 is shown. The tip shroud 50 of FIG. 7 includes a single cutter member 90 disposed proximate a central location between the first edge 60 and the second edge 62 of the tip blade 52. The cutter member 90 extends away from both the first surface 56 and the second surface 58, such that regardless of the deformation of the tip shroud 50, or tip blade 52, during operation, the cutter member 90 may function to remove material from the aforementioned counterpart surface, such as a honeycomb structure. In the illustrated embodiment, the substrate 70 is flared outward from the first surface 56 at a location proximate the second edge 62. Similarly, an additional substrate portion may be flared outward at various other locations, such as away from the second surface 58 at a location proximate the first edge 60.
  • While the invention has been described in detail in connection with only a limited number of embodiments, it should be readily understood that the invention is not limited to such disclosed embodiments. Rather, the invention can be modified to incorporate any number of variations, alterations, substitutions or equivalent arrangements not heretofore described, but which are commensurate with the spirit and scope of the invention. Additionally, while various embodiments of the invention have been described, it is to be understood that aspects of the invention may include only some of the described embodiments. Accordingly, the invention is not to be seen as limited by the foregoing description, but is only limited by the scope of the appended claims.

Claims (20)

1. A rotating assembly for a turbine assembly comprising:
an airfoil extending radially outward from, and rotatable about, an axial centerline;
a tip shroud integrally connected proximate a radially outer tip of the airfoil;
a substrate operably coupled to the tip shroud; and
at least one hard face interface member secured to the substrate.
2. The rotating assembly of claim 1, the tip shroud comprising:
a tip blade having a first surface, a second surface opposing the first surface, a first edge and a second edge,
a first cutter member disposed proximate the first edge and the second surface; and
a second cutter member disposed proximate the second edge and the first surface.
3. The rotating assembly of claim 1, the tip shroud comprising a relatively planar tip blade having a first edge and a second edge, wherein the substrate includes a fillet portion proximate at least one of the first edge or the second edge.
4. The rotating assembly of claim 1, wherein the at least one hard face interface member includes a base portion and a plurality of edge portions, wherein a portion of the substrate at least partially surrounds at least one of the plurality of edge portions.
5. The rotating assembly of claim 1, wherein the at least one hard face interface member is formed from a pre-sintered preformed material.
6. The rotating assembly of claim 1, wherein the substrate includes a plurality of grooved portions.
7. The rotating assembly of claim 6, wherein the plurality of grooved portions are relatively U-shaped.
8. The rotating assembly of claim 6, wherein the plurality of grooved portions are relatively V-shaped.
9. The rotating assembly of claim 1, wherein the substrate includes a plurality of protrusions.
10. The rotating assembly of claim 1, wherein the at least one hard face interface member is brazed to the substrate.
11. A rotating assembly for a turbine assembly comprising:
a plurality of rotating members extending radially outward from, and rotatable about, an axial centerline;
a tip shroud integrally connected proximate a radially outer tip of at least one of the plurality of rotating members;
a substrate operably coupled to the tip shroud; and
a hard face interface member having a base portion and a plurality of edge portions, wherein the base portion is secured to the substrate and a portion of the substrate at least partially surrounds at least one of the plurality of edge portions.
12. The rotating assembly of claim 11, the tip shroud comprising:
a tip blade having a first surface, a second surface opposing the first surface, a first edge and a second edge,
a first cutter member disposed proximate the first edge and the second surface; and
a second cutter member disposed proximate the second edge and the first surface.
13. The rotating assembly of claim 11, the tip shroud comprising a relatively planar tip blade having a first edge and a second edge, wherein the substrate includes a fillet portion proximate at least one of the first edge or the second edge.
14. The rotating assembly of claim 11, wherein the hard face interface member is formed from a pre-sintered preformed material.
15. The rotating assembly of claim 11, wherein the substrate includes a plurality of grooved portions.
16. The rotating assembly of claim 11, wherein the substrate includes a plurality of protrusions.
17. A rotating assembly for a turbine assembly comprising:
an airfoil extending radially outward from, and rotatable about, an axial centerline;
a tip shroud integrally connected proximate a radially outer tip of the airfoil;
a substrate operably coupled to the tip shroud, wherein the substrate includes a plurality of grooved portions; and
at least one hard face interface member being both mechanically interlocked with, and metallurgically bonded to the substrate.
18. The rotating assembly of claim 17, wherein the plurality of grooved portions are relatively U-shaped.
19. The rotating assembly of claim 17, wherein the plurality of grooved portions are relatively V-shaped.
20. The rotating assembly of claim 17, wherein the at least one hard face interface member is formed from a pre-sintered preformed material.
US13/368,931 2012-02-08 2012-02-08 Rotating assembly for a turbine assembly Abandoned US20130202439A1 (en)

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Application Number Priority Date Filing Date Title
US13/368,931 US20130202439A1 (en) 2012-02-08 2012-02-08 Rotating assembly for a turbine assembly

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Application Number Priority Date Filing Date Title
US13/368,931 US20130202439A1 (en) 2012-02-08 2012-02-08 Rotating assembly for a turbine assembly
JP2013016279A JP6239237B2 (en) 2012-02-08 2013-01-31 Rotating assembly for turbine assembly
EP13154014.8A EP2626517B1 (en) 2012-02-08 2013-02-05 Rotating assembly for a turbine assembly

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US20130202439A1 true US20130202439A1 (en) 2013-08-08

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Cited By (1)

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Publication number Priority date Publication date Assignee Title
WO2017155497A1 (en) * 2016-03-07 2017-09-14 Siemens Aktiengesellschaft Gas turbine blade tip shroud sealing and flow guiding features

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JP6567394B2 (en) * 2015-11-10 2019-08-28 株式会社東芝 Method for repairing worn parts of gas turbine parts
DE102019202387A1 (en) * 2019-02-21 2020-08-27 MTU Aero Engines AG Blade for a high-speed turbine stage with a single sealing element

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US6224963B1 (en) * 1997-05-14 2001-05-01 Alliedsignal Inc. Laser segmented thick thermal barrier coatings for turbine shrouds
US20050079058A1 (en) * 2003-10-09 2005-04-14 Pratt & Whitney Canada Corp. Shrouded turbine blades with locally increased contact faces
US20050191182A1 (en) * 2004-02-26 2005-09-01 Richard Seleski Turbine blade shroud cutter tip
US20080292466A1 (en) * 2007-05-24 2008-11-27 General Electric Company Method to center locate cutter teeth on shrouded turbine blades
US20090202344A1 (en) * 2008-02-13 2009-08-13 General Electric Company Rotating assembly for a turbomachine
US7771171B2 (en) * 2006-12-14 2010-08-10 General Electric Company Systems for preventing wear on turbine blade tip shrouds

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US4764089A (en) * 1986-08-07 1988-08-16 Allied-Signal Inc. Abradable strain-tolerant ceramic coated turbine shroud
US6164916A (en) * 1998-11-02 2000-12-26 General Electric Company Method of applying wear-resistant materials to turbine blades, and turbine blades having wear-resistant materials
US7686568B2 (en) * 2006-09-22 2010-03-30 General Electric Company Methods and apparatus for fabricating turbine engines
ES2580227T3 (en) * 2009-07-22 2016-08-22 MTU Aero Engines AG Method for coating turbine blades

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Publication number Priority date Publication date Assignee Title
US6224963B1 (en) * 1997-05-14 2001-05-01 Alliedsignal Inc. Laser segmented thick thermal barrier coatings for turbine shrouds
US20050079058A1 (en) * 2003-10-09 2005-04-14 Pratt & Whitney Canada Corp. Shrouded turbine blades with locally increased contact faces
US20050191182A1 (en) * 2004-02-26 2005-09-01 Richard Seleski Turbine blade shroud cutter tip
US7771171B2 (en) * 2006-12-14 2010-08-10 General Electric Company Systems for preventing wear on turbine blade tip shrouds
US20080292466A1 (en) * 2007-05-24 2008-11-27 General Electric Company Method to center locate cutter teeth on shrouded turbine blades
US20090202344A1 (en) * 2008-02-13 2009-08-13 General Electric Company Rotating assembly for a turbomachine

Cited By (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
WO2017155497A1 (en) * 2016-03-07 2017-09-14 Siemens Aktiengesellschaft Gas turbine blade tip shroud sealing and flow guiding features

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Publication number Publication date
JP6239237B2 (en) 2017-11-29
EP2626517B1 (en) 2020-04-01
JP2013160228A (en) 2013-08-19
EP2626517A3 (en) 2017-03-29
EP2626517A2 (en) 2013-08-14

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