US20130156966A1 - Method for reprocessing a turbine blade having at least one platform - Google Patents

Method for reprocessing a turbine blade having at least one platform Download PDF

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Publication number
US20130156966A1
US20130156966A1 US13/634,642 US201113634642A US2013156966A1 US 20130156966 A1 US20130156966 A1 US 20130156966A1 US 201113634642 A US201113634642 A US 201113634642A US 2013156966 A1 US2013156966 A1 US 2013156966A1
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United States
Prior art keywords
platform
vane
adhesion promoter
turbine blade
blade
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Abandoned
Application number
US13/634,642
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English (en)
Inventor
Martin Grohnert
Andreas Oppert
Gerhard Reich
Rolf Wilkenhöner
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Siemens AG
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Siemens AG
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Filing date
Publication date
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Assigned to SIEMENS AKTIENGESELLSCHAFT reassignment SIEMENS AKTIENGESELLSCHAFT ASSIGNMENT OF ASSIGNORS INTEREST (SEE DOCUMENT FOR DETAILS). Assignors: OPPERT, ANDREAS, GROHNERT, MARTIN, WILKENHOENER, ROLF, REICH, GERHARD
Publication of US20130156966A1 publication Critical patent/US20130156966A1/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
    • F01D5/005—Repairing methods or devices
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B23—MACHINE TOOLS; METAL-WORKING NOT OTHERWISE PROVIDED FOR
    • B23P—METAL-WORKING NOT OTHERWISE PROVIDED FOR; COMBINED OPERATIONS; UNIVERSAL MACHINE TOOLS
    • B23P6/00—Restoring or reconditioning objects
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B23—MACHINE TOOLS; METAL-WORKING NOT OTHERWISE PROVIDED FOR
    • B23P—METAL-WORKING NOT OTHERWISE PROVIDED FOR; COMBINED OPERATIONS; UNIVERSAL MACHINE TOOLS
    • B23P6/00—Restoring or reconditioning objects
    • B23P6/002—Repairing turbine components, e.g. moving or stationary blades, rotors
    • B23P6/007—Repairing turbine components, e.g. moving or stationary blades, rotors using only additive methods, e.g. build-up welding
    • 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
    • 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
    • F05D2230/00—Manufacture
    • F05D2230/30—Manufacture with deposition of material
    • 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
    • F05D2230/00—Manufacture
    • F05D2230/30—Manufacture with deposition of material
    • F05D2230/31—Layer deposition
    • 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
    • F05D2240/00—Components
    • F05D2240/80—Platforms for stationary or moving 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 relates to a process for refurbishing a turbine blade or vane having at least one platform, wherein the turbine blade or vane can be formed in particular as a gas turbine blade or vane.
  • a liquid or gaseous fuel is burned in a combustion chamber and the hot gases under high pressure which form during the combustion are fed to the turbine, where they transfer momentum to the rotor blades of a turbine with expansion and cooling.
  • the transfer of momentum to the rotor blades is optimized by means of guide vanes.
  • the turbine blades or vanes are produced from superalloys which can withstand high temperatures, and are additionally coated with a thermal barrier coating system, in order to further increase the resistance of the blades or vanes to the oxidizing and corrosive conditions in the hot gas.
  • a coating typically comprises a ceramic thermal barrier coating which is bonded to the superalloy material of the blade or vane by means of an adhesion promoter layer.
  • Typical adhesion promoter layers are so-called MCrAlX layers, in which M stands for iron (Fe), cobalt (Co), nickel (Ni) or a combination of these metals.
  • M stands for iron (Fe), cobalt (Co), nickel (Ni) or a combination of these metals.
  • X represents an active element and stands for yttrium (Y) and/or silicon (Si) and/or at least one rare earth element or hafnium (Hf).
  • Such alloys are known, for example, from EP 0 486 489 B1, EP 0 786 017 B1, EP 0 412 397 B1 or EP 1 306 454 A1.
  • EP 1 808 266 A2 proposes removing platform regions damaged by corrosion in the region of the trailing edge of the turbine blade or vane and then rebuilding the removed region by build-up welding and subsequent grinding to the correct dimension.
  • the undersized side faces of platforms can also be built up again in this way, build-up welding on superalloy materials is difficult.
  • undesirable structural properties in the superalloy material which weaken the material can arise on account of the introduction of heat.
  • the desired dimension of the platform is restored by applying material to the at least one platform side face in such a manner that, after the material application, the platform is oversized, and then the platform is given the desired dimension by machining the at least one platform side face.
  • the material application is effected with the material of an adhesion promoter layer.
  • This material can be, in particular, an MCrAlX material.
  • the invention has the advantage that the application of an adhesion promoter material, in particular of MCrAlX material, does not entail such a high introduction of heat into the superalloy material as would be the case, for example, in the case of build-up welding.
  • an adhesion promoter material in particular of MCrAlX material
  • the material application can be integrated in the process of recoating the turbine blade or vane, since an adhesion promoter layer is also applied when reapplying a thermal barrier coating system.
  • the process according to the invention therefore makes it possible, in a cost-effective and gentle manner, to restore the desired dimension of platform side faces in operationally stressed turbine blades or vanes, as a result of which the proportion of rejects of operationally stressed turbine blades or vanes can be reduced.
  • the material application can be effected in particular by means of the repeated application of adhesion promoter material.
  • a material application of at least 10 ⁇ m, preferably at least 30 ⁇ m, can be effected in this case in particular each time the application of adhesion promoter material is repeated.
  • a bonding heat treatment can take place after the application of the adhesion promoter material.
  • the material application and the machining can also be effected in particular on two oppositely positioned platform faces of a blade or vane platform. Specifically, it is often the case that undersized regions caused by corrosion arise on two oppositely positioned sides of blade or vane platforms at the same time.
  • the turbine blade or vane has a central axis. It is therefore advantageous if the machining of the platform side faces to give the platform the desired dimension again after the application of material is effected with respect to the central axis. It is thereby possible to ensure that not only the platform width but also the distance between the platform side faces and the main blade or vane part of the turbine are given the desired dimension again.
  • the current dimension of the platform can be gathered by scanning at least five measurement points on the opposite platform side faces. The required material removal for the machining is then determined from the current dimension.
  • the two opposite platform faces can in particular be the platform side faces which are located on the pressure side and suction side in relation to a main blade or vane part with a pressure side and a suction side.
  • these side faces are typically exposed to the hot gas oxidation and the resulting corrosion to a greater extent than the platform side faces located on the inflow side and on the outflow side.
  • the machining can be realized in particular by face grinding.
  • the process can comprise the removal of layers from the turbine blade or vane before the thermal barrier coating system is renewed.
  • activation blasting can be effected after the layers have been removed and before the thermal barrier coating system is renewed.
  • the activation blasting would then also include in particular the platform side faces onto which material is to be applied.
  • the surfaces are irradiated by means of a blasting agent, for example by means of aluminum oxide (Al 2 O 3 ), as a result of which the surface is roughened, which improves the adhesion of the adhesion promoter material to be applied.
  • the adhesion promoter material can be applied using a thermal spraying process, for example plasma spraying, flame spraying, etc.
  • a thermal spraying process for example plasma spraying, flame spraying, etc.
  • Such processes are known as possible processes for applying adhesion promoter layers and can therefore also be used in a readily manageable manner for the application of material to undersized platform side faces.
  • FIG. 1 shows a schematic illustration of a gas turbine in a partial longitudinal section.
  • FIG. 2 shows an example of a combustion chamber of a gas turbine in a partially sectioned, perspective illustration.
  • FIG. 3 shows an example of a turbine blade or vane in a perspective illustration.
  • FIG. 4 shows a schematic plan view of a turbine blade or vane which is undersized as a result of corrosion on side faces of the platform.
  • FIG. 5 shows the turbine blade or vane shown in FIG. 4 during the application of adhesion promoter material.
  • FIG. 6 shows the turbine blade or vane shown in FIG. 4 during the grinding of the applied adhesion promoter material to the desired dimension.
  • FIG. 1 shows, by way of example, a partial longitudinal section through a gas turbine 100 .
  • the gas turbine 100 has a rotor 103 with a shaft 101 which is mounted such that it can rotate about an axis of rotation 102 and is also referred to as the turbine rotor.
  • the annular combustion chamber 110 is in communication with a, for example, annular hot-gas passage 111 , where, by way of example, four successive turbine stages 112 form the turbine 108 .
  • Each turbine stage 112 is formed, for example, from two blade or vane rings. As seen in the direction of flow of a working medium 113 , in the hot-gas passage 111 a row of guide vanes 115 is followed by a row 125 formed from rotor blades 120 .
  • the compressor 105 While the gas turbine 100 is operating, the compressor 105 sucks in air 135 through the intake housing 104 and compresses it. The compressed air provided at the turbine-side end of the compressor 105 is passed to the burners 107 , where it is mixed with a fuel. The mix is then burnt in the combustion chamber 110 , forming the working medium 113 . From there, the working medium 113 flows along the hot-gas passage 111 past the guide vanes 130 and the rotor blades 120 . The working medium 113 is expanded at the rotor blades 120 , transferring its momentum, so that the rotor blades 120 drive the rotor 103 and the latter in turn drives the generator coupled to it.
  • iron-based, nickel-based or cobalt-based superalloys are used as material for the components, in particular for the turbine blade or vane 120 , 130 and components of the combustion chamber 110 .
  • the blades or vanes 120 , 130 may likewise have coatings protecting against corrosion (MCrAlX; M is at least one element selected from the group consisting of iron (Fe), cobalt (Co), nickel (Ni), X is an active element and stands for yttrium (Y) and/or silicon, scandium (Sc) and/or at least one rare earth element, or hafnium). Alloys of this type are known from EP 0 486 489 B1, EP 0 786 017 B1, EP 0 412 397 B1 or EP 1 306 454 A1.
  • Columnar grains are produced in the thermal barrier coating by suitable coating processes, such as for example electron beam physical vapor deposition (EB-PVD).
  • EB-PVD electron beam physical vapor deposition
  • the guide vane 130 has a guide vane root (not shown here), which faces the inner housing 138 of the turbine 108 , and a guide vane head which is at the opposite end from the guide vane root.
  • the guide vane head faces the rotor 103 and is fixed to a securing ring 140 of the stator 143 .
  • FIG. 2 shows a combustion chamber 110 of a gas turbine.
  • the combustion chamber 110 is designed for a relatively high temperature of the working medium M of approximately 1000° C. to 1600° C.
  • the combustion chamber wall 153 is provided, on its side which faces the working medium M, with an inner lining formed from heat shield elements 155 .
  • M is at least one element selected from the group consisting of iron (Fe), cobalt (Co), nickel (Ni), X is an active element and stands for yttrium (Y) and/or silicon and/or at least one rare earth element or hafnium (Hf). Alloys of this type are known from EP 0 486 489 B1, EP 0 786 017 B1, EP 0 412 397 B1 or EP 1 306 454 A1.
  • Columnar grains are produced in the thermal barrier coating by suitable coating processes, such as for example electron beam physical vapor deposition (EB-PVD).
  • EB-PVD electron beam physical vapor deposition
  • Refurbishment means that after they have been used, protective layers may have to be removed from heat shield elements 155 (e.g. by sand-blasting). Then, the corrosion and/or oxidation layers and products are removed. If appropriate, cracks in the heat shield element 155 are also repaired. This is followed by recoating of the heat shield elements 155 , after which the heat shield elements 155 can be reused.
  • a cooling system may be provided for the heat shield elements 155 and/or their holding elements, on account of the high temperatures in the interior of the combustion chamber 110 .
  • the heat shield elements 155 are then, for example, hollow and may also have cooling holes (not shown) opening out into the combustion chamber space 154 .
  • the turbomachine may be a gas turbine of an aircraft or of a power plant for generating electricity, a steam turbine or a compressor.
  • the blade or vane 120 , 130 has, in succession along the longitudinal axis 121 , a securing region 400 , an adjoining blade or vane platform 403 and a main blade or vane part 406 and a blade or vane tip 415 .
  • the vane 130 may have a further platform (not shown) at its vane tip 415 .
  • a blade or vane root 183 which is used to secure the rotor blades 120 , 130 to a shaft or a disk (not shown), is formed in the securing region 400 .
  • the blade or vane 120 , 130 has a leading edge 409 and a trailing edge 412 for a medium which flows past the main blade or vane part 406 .
  • the blade or vane 120 , 130 may in this case be produced by a casting process, by means of directional solidification, by a forging process, by a milling process or combinations thereof.
  • Workpieces with a single-crystal structure or structures are used as components for machines which, in operation, are exposed to high mechanical, thermal and/or chemical stresses.
  • Single-crystal workpieces of this type are produced, for example, by directional solidification from the melt. This involves casting processes in which the liquid metallic alloy solidifies to form the single-crystal structure, i.e. the single-crystal workpiece, or solidifies directionally.
  • dendritic crystals are oriented along the direction of heat flow and form either a columnar crystalline grain structure (i.e. grains which run over the entire length of the workpiece and are referred to here, in accordance with the language customarily used, as directionally solidified) or a single-crystal structure, i.e. the entire workpiece consists of one single crystal.
  • a transition to globular (polycrystalline) solidification needs to be avoided, since non-directional growth inevitably forms transverse and longitudinal grain boundaries, which negate the favorable properties of the directionally solidified or single-crystal component.
  • directionally solidified microstructures refers in general terms to directionally solidified microstructures, this is to be understood as meaning both single crystals, which do not have any grain boundaries or at most have small-angle grain boundaries, and columnar crystal structures, which do have grain boundaries running in the longitudinal direction but do not have any transverse grain boundaries.
  • This second form of crystalline structures is also described as directionally solidified microstructures (directionally solidified structures).
  • the blades or vanes 120 , 130 may likewise have coatings protecting against corrosion or oxidation e.g. (MCrAlX; M is at least one element selected from the group consisting of iron (Fe), cobalt (Co), nickel (Ni), X is an active element and stands for yttrium (Y) and/or silicon and/or at least one rare earth element, or hafnium (Hf). Alloys of this type are known from EP 0 486 489 B1, EP 0 786 017 B1, EP 0 412 397 B1 or EP 1 306 454 A1.
  • MrAlX M is at least one element selected from the group consisting of iron (Fe), cobalt (Co), nickel (Ni)
  • X is an active element and stands for yttrium (Y) and/or silicon and/or at least one rare earth element, or hafnium (Hf). Alloys of this type are known from EP 0 486 489 B1, EP 0 786 017
  • the density is preferably 95% of the theoretical density.
  • the layer preferably has a composition Co-30Ni-28Cr-8Al-0.6Y-0.7Si or Co-28Ni-24Cr-10Al-0.6Y.
  • nickel-based protective layers such as Ni-10Cr-12Al-0.6Y-3Re or Ni-12Co-21Cr-11Al-0.4Y-2Re or Ni-25Co-17Cr-10Al-0.4Y-1.5Re.
  • thermal barrier coating which is preferably the outermost layer, to be present on the MCrAlX, consisting for example of ZrO 2 , Y 2 O 3 —ZrO 2 , i.e. unstabilized, partially stabilized or fully stabilized by yttrium oxide and/or calcium oxide and/or magnesium oxide.
  • the thermal barrier coating covers the entire MCrAlX layer.
  • Columnar grains are produced in the thermal barrier coating by suitable coating processes, such as for example electron beam physical vapor deposition (EB-PVD).
  • EB-PVD electron beam physical vapor deposition
  • the thermal barrier coating may include grains that are porous or have micro-cracks or macro-cracks, in order to improve the resistance to thermal shocks.
  • the thermal barrier coating is therefore preferably more porous than the MCrAlX layer.
  • Refurbishment means that after they have been used, protective layers may have to be removed from components 120 , 130 (e.g. by sand-blasting). Then, the corrosion and/or oxidation layers and products are removed. If appropriate, cracks in the component 120 , 130 are also repaired. This is followed by recoating of the component 120 , 130 , after which the component 120 , 130 can be reused.
  • the blade or vane 120 , 130 may be hollow or solid in form. If the blade or vane 120 , 130 is to be cooled, it is hollow and may also have film-cooling holes 418 (indicated by dashed lines).
  • FIG. 4 shows a schematic illustration of a corroded turbine blade or vane.
  • the illustration shows the main blade or vane part 1 with a suction side 3 and pressure side 5 and also the blade or vane platform 7 .
  • said blade or vane platform is undersized both on the suction-side platform side face 9 and on the pressure-side platform side face 11 .
  • the desired dimension of the platform 7 is shown by dashed lines in the figure. At this point, it should be pointed out that the undersize induced by corrosion is shown in exaggerated form in order to enhance the clarity of the drawings.
  • Corrosion in regions of the turbine blade or vane other than the side faces 9 , 11 of the blade or vane platform 7 is typically also present, but is not shown in the figures so as not to unnecessarily complicate the drawings.
  • the turbine blade or vane which has been corroded by the harsh ambient conditions which prevail during gas turbine operation is subjected according to the invention to refurbishment, in which the desired dimension of the blade or vane platform, in particular at the corroded side faces 9 , 11 , is restored.
  • the restoration of the desired dimension is in part integrated in the procedure for applying a new thermal barrier coating system to the turbine blade or vane.
  • the old coating is firstly removed from the turbine blade or vane, for example by means of suitable solutions and/or suitable blasting processes, and the blade or vane is then cleaned in order to remove possible oxidation residues.
  • activation blasting for example by means of aluminum oxide particles (Al 2 O 3 , corundum), by means of which the surface is roughened.
  • the thus prepared turbine blade or vane is then introduced into a coating apparatus in order to then coat the main blade or vane part 1 and the face 13 of the blade or vane platform 7 which faces toward the hot loading path with the thermal barrier coating system.
  • the side faces of the blade or vane platform 7 are usually masked or obscured in the prior art, since they are not intended to be provided with a thermal barrier coating system.
  • the suction-side platform side face 9 and also the pressure-side platform side face 11 are not masked, or obscured, but instead are left free. It is thereby possible to apply coating material to these two faces.
  • the coating process firstly involves the application of an adhesion promoter layer, which in the present example is in the form of an MCrAlX layer.
  • the adhesion promoter layer is applied by means of a thermal spraying process, for example by means of plasma spraying or flame spraying.
  • the MCrAlX material is applied not only to the main blade or vane part 1 and the top side 13 of the blade or vane platform 7 , but also to the suction-side platform side face 9 and the pressure-side platform side face 11 .
  • the spraying process is indicated in FIG. 5 by a schematically shown spray nozzle 15 .
  • a plurality of layers of the MCrAlX material are applied to the suction-side platform side face 9 and the pressure-side platform side face 11 , with each layer having a minimum thickness of 10 ⁇ m, preferably of 30 ⁇ m.
  • the layered application of MCrAlX material 12 to the platform side faces 9 , 11 is effected until the desired dimension of the platform 7 , as indicated by dashed lines in FIGS. 4 to 6 , is exceeded.
  • This state is shown in FIG. 5 for the suction-side platform side face 9
  • the layered spraying-on of the MCrAlX material 12 by means of the spray nozzle 15 is shown for the pressure-side platform side face 11 .
  • a preferred bonding heat treatment which improves the bonding of the applied MCrAlX material 12 to the superalloy material of the turbine blade or vane is carried out.
  • a thermal barrier coating for example a zirconium oxide layer (ZrO 2 ), the structure of which is stabilized at least partially by yttrium oxide (Y 2 O 3 ), to the MCrAlX layer.
  • the thermal barrier coating is applied in particular to the main blade or vane part 1 and the surface 13 of the platform 7 .
  • thermal barrier coating is also applied to the MCrAlX material 12 applied to the platform side faces 9 , 11 .
  • the thermal barrier coating can likewise be applied by means of a thermal spraying process. Alternatively, however, it is also possible to produce the thermal barrier coating by vapor deposition.
  • the turbine blade or vane is removed from the coating apparatus and clamped in an apparatus for machining, in which, in the present exemplary embodiment, the now oversized suction-side and pressure-side platform side faces are then machined by means of grinding.
  • the turbine blade or vane is clamped in such a way here that the central axis A of the turbine blade or vane is the same as the central axis of the clamping apparatus.
  • the turbine blade or vane is clamped for grinding such that the turbine blade or vane can be freely positioned in space with respect to its central axis A, i.e. can be rotated through 360°.
  • the current width b of the blade or vane platform 7 is then captured by sampling at least 5 measurement points for each platform side face 9 , 11 on the pressure side and suction side.
  • a computer program is then used to calculate the material removal which is required to give the blade or vane platform 7 having the applied MCrAlX material the desired dimension.
  • the machining removal calculated for the suction-side platform side face 9 and the pressure-side platform side face 11 is in this case based on the blade or vane central axis A.
  • the ascertained material to be removed is then machined off by face grinding by means of a grinding apparatus, which is shown in greatly schematic form in FIG. 6 under the reference numeral 17 . Before zero grinding, the grinding disk is removed and the truing amount is compensated by the program, based on the disk diameter of the grinding apparatus. After the grinding process has been completed, the width b of the blade or vane platform 7 has the desired dimension again.

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  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
  • Coating By Spraying Or Casting (AREA)
US13/634,642 2010-03-19 2011-03-15 Method for reprocessing a turbine blade having at least one platform Abandoned US20130156966A1 (en)

Applications Claiming Priority (3)

Application Number Priority Date Filing Date Title
EP10002967A EP2366488A1 (de) 2010-03-19 2010-03-19 Verfahren zum Wiederaufarbeiten einer Turbinenschaufel mit wenigstens einer Plattform
EP10002967.7 2010-03-19
PCT/EP2011/053899 WO2011113833A1 (de) 2010-03-19 2011-03-15 Verfahren zum wiederaufarbeiten einer turbinenschaufel mit wenigstens einer plattform

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US20130156966A1 true US20130156966A1 (en) 2013-06-20

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US (1) US20130156966A1 (de)
EP (2) EP2366488A1 (de)
JP (1) JP2013522526A (de)
KR (2) KR20140119820A (de)
CN (1) CN102811835A (de)
RU (1) RU2527509C2 (de)
WO (1) WO2011113833A1 (de)

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US9638051B2 (en) 2013-09-04 2017-05-02 General Electric Company Turbomachine bucket having angel wing for differently sized discouragers and related methods

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US20130318787A1 (en) * 2012-06-05 2013-12-05 Seth J. Thomen Manufacturing a family of airfoils
DE102012013577A1 (de) 2012-07-10 2014-01-16 Oerlikon Trading Ag, Trübbach Hochleistungsimpulsbeschichtungsmethode
EP2918783A1 (de) * 2014-03-12 2015-09-16 Siemens Aktiengesellschaft Turbinenschaufel mit einer beschichteten Plattform
DE102014224865A1 (de) * 2014-12-04 2016-06-09 Siemens Aktiengesellschaft Verfahren zur Beschichtung einer Turbinenschaufel
RU2619374C1 (ru) * 2016-01-29 2017-05-15 Арсений Евгеньевич Ляшенко Способ удаления царапин и сколов с лакокрасочного покрытия автомобиля

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WO2011113833A1 (de) 2011-09-22
KR20140119820A (ko) 2014-10-10
EP2366488A1 (de) 2011-09-21
EP2547488B1 (de) 2018-08-29
CN102811835A (zh) 2012-12-05
RU2527509C2 (ru) 2014-09-10
RU2012144432A (ru) 2014-04-27

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