EP2589682A1 - Couche d'isolation thermique en céramique sur une surface structurée et procédé de fabrication - Google Patents

Couche d'isolation thermique en céramique sur une surface structurée et procédé de fabrication Download PDF

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Publication number
EP2589682A1
EP2589682A1 EP11188032.4A EP11188032A EP2589682A1 EP 2589682 A1 EP2589682 A1 EP 2589682A1 EP 11188032 A EP11188032 A EP 11188032A EP 2589682 A1 EP2589682 A1 EP 2589682A1
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EP
European Patent Office
Prior art keywords
layer
recesses
layer system
turbine
structured surface
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.)
Withdrawn
Application number
EP11188032.4A
Other languages
German (de)
English (en)
Inventor
Fathi Ahmad
Christian Amann
Björn Beckmann
Björn Buchholz
Giuseppe Gaio
Thomas Hille
Eckart Schumann
Rostislav Teteruk
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.)
Siemens AG
Original Assignee
Siemens AG
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 Siemens AG filed Critical Siemens AG
Priority to EP11188032.4A priority Critical patent/EP2589682A1/fr
Priority to US14/354,573 priority patent/US9862002B2/en
Priority to EP12759691.4A priority patent/EP2753729A1/fr
Priority to PCT/EP2012/068048 priority patent/WO2013068159A1/fr
Publication of EP2589682A1 publication Critical patent/EP2589682A1/fr
Withdrawn legal-status Critical Current

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Classifications

    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B05—SPRAYING OR ATOMISING IN GENERAL; APPLYING FLUENT MATERIALS TO SURFACES, IN GENERAL
    • B05D—PROCESSES FOR APPLYING FLUENT MATERIALS TO SURFACES, IN GENERAL
    • B05D3/00—Pretreatment of surfaces to which liquids or other fluent materials are to be applied; After-treatment of applied coatings, e.g. intermediate treating of an applied coating preparatory to subsequent applications of liquids or other fluent materials
    • B05D3/007—After-treatment
    • C—CHEMISTRY; METALLURGY
    • C23—COATING METALLIC MATERIAL; COATING MATERIAL WITH METALLIC MATERIAL; CHEMICAL SURFACE TREATMENT; DIFFUSION TREATMENT OF METALLIC MATERIAL; COATING BY VACUUM EVAPORATION, BY SPUTTERING, BY ION IMPLANTATION OR BY CHEMICAL VAPOUR DEPOSITION, IN GENERAL; INHIBITING CORROSION OF METALLIC MATERIAL OR INCRUSTATION IN GENERAL
    • C23C—COATING METALLIC MATERIAL; COATING MATERIAL WITH METALLIC MATERIAL; SURFACE TREATMENT OF METALLIC MATERIAL BY DIFFUSION INTO THE SURFACE, BY CHEMICAL CONVERSION OR SUBSTITUTION; COATING BY VACUUM EVAPORATION, BY SPUTTERING, BY ION IMPLANTATION OR BY CHEMICAL VAPOUR DEPOSITION, IN GENERAL
    • C23C30/00—Coating with metallic material characterised only by the composition of the metallic material, i.e. not characterised by the coating process
    • 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
    • F01D5/28—Selecting particular materials; Particular measures relating thereto; Measures against erosion or corrosion
    • 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
    • F01D5/28—Selecting particular materials; Particular measures relating thereto; Measures against erosion or corrosion
    • F01D5/288—Protective coatings for blades
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
    • F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
    • F23R3/007—Continuous combustion chambers using liquid or gaseous fuel constructed mainly of ceramic components
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2250/00—Geometry
    • F05D2250/20—Three-dimensional
    • F05D2250/29—Three-dimensional machined; miscellaneous
    • F05D2250/294—Three-dimensional machined; miscellaneous grooved
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23M—CASINGS, LININGS, WALLS OR DOORS SPECIALLY ADAPTED FOR COMBUSTION CHAMBERS, e.g. FIREBRIDGES; DEVICES FOR DEFLECTING AIR, FLAMES OR COMBUSTION PRODUCTS IN COMBUSTION CHAMBERS; SAFETY ARRANGEMENTS SPECIALLY ADAPTED FOR COMBUSTION APPARATUS; DETAILS OF COMBUSTION CHAMBERS, NOT OTHERWISE PROVIDED FOR
    • F23M2900/00—Special features of, or arrangements for combustion chambers
    • F23M2900/05003—Details of manufacturing specially adapted for combustion chambers
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23M—CASINGS, LININGS, WALLS OR DOORS SPECIALLY ADAPTED FOR COMBUSTION CHAMBERS, e.g. FIREBRIDGES; DEVICES FOR DEFLECTING AIR, FLAMES OR COMBUSTION PRODUCTS IN COMBUSTION CHAMBERS; SAFETY ARRANGEMENTS SPECIALLY ADAPTED FOR COMBUSTION APPARATUS; DETAILS OF COMBUSTION CHAMBERS, NOT OTHERWISE PROVIDED FOR
    • F23M2900/00—Special features of, or arrangements for combustion chambers
    • F23M2900/05004—Special materials for walls or lining
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
    • F23R2900/00—Special features of, or arrangements for continuous combustion chambers; Combustion processes therefor
    • F23R2900/00018—Manufacturing combustion chamber liners or subparts

Definitions

  • the invention relates to a ceramic layer on a structured surface and manufacturing method.
  • High-temperature components such as gas turbine components are often provided with ceramic thermal barrier coatings, which can also chip off under the most extreme operating conditions. This happens because stresses occur that lead to flaking of the ceramic thermal barrier coating.
  • the object is achieved by a ceramic thermal barrier coating according to claim 1 and a manufacturing method according to claim 9 or 12.
  • FIG. 5 a layer system 1, 120, 130, 155 is shown.
  • the layer system 1, 120, 130, 155 has a substrate 4 which has in particular a nickel- or cobalt-based superalloy, in particular consists thereof, very particularly according to an alloy according to FIG. 9 ,
  • an intermediate layer 10 in particular a metallic adhesion promoter layer 10 is optionally present, on whose surface 13, in turn, a ceramic thermal barrier coating 16 is present.
  • a ceramic thermal barrier coating 16 is present.
  • the metallic adhesion promoter layer 10 preferably comprises an MCrAIX alloy.
  • depressions 19 ', 19 ",... are present or are introduced into the surface 7 of the substrate 4 or in the surface 13 of the layer 10 (FIG. Fig. 1 ).
  • the recesses 19 ', 19'', ... have a certain depth b and a certain width a.
  • the width a of the recesses 19 ', 19 ", ... is at least 10 .mu.m, preferably 10 .mu.m to 30 .mu.m.
  • the depth b is at least 10%, preferably 10% to 30% of the thickness of the underlying layer 10, more preferably 10 ⁇ m to 30 ⁇ m.
  • the distance d of the opposite recesses 19 ', 19 ", ... is at least 100 ⁇ m, preferably between 100 ⁇ m and 300 ⁇ m ( Fig. 2 ).
  • the parameters a, b, d can be varied on the surface 7, 13 depending on the conditions of use or locally (on the blade leaf 406, but not on the blade blade shape 403).
  • the recesses 19 ', 19 "only locally limited on the surface 7, 13 of the component 1, 120, 130 may be present.
  • the depressions 19 ', 19 ",... can preferably be made round on the base 20 (FIG. Fig. 1 ).
  • the recesses 19 ', 19 ", ... can be a honeycomb structure ( FIG. 3 ) or a mesh structure ( FIG. 4 ) exhibit.
  • FIG. 1 In FIG. 1 is shown a cross section through such a deliberately structured surface. Depending on how large the depressions 19 ', 19 “,..., The depression 19', 19” also continues on the surface 22 of the ceramic thermal barrier coating 16 in depressions 23 ', 23 ".
  • the coating 16 may be configured so that the outermost surface 22 is smooth, i. the underlying recesses 23 ', 23 "would not be visible on the surface 22.
  • the layers 10 are applied by applying material (eg powder) from a nozzle, in particular linear.
  • material eg powder
  • the structured surface 7, 13 is an integral part of a layer 10. It thus does not represent a honeycomb structure which is filled with a ceramic material.
  • FIG. 6 shows by way of example a gas turbine 100 in a longitudinal partial section.
  • the gas turbine 100 has inside a rotatably mounted about a rotation axis 102 rotor 103 with a shaft 101, which is also referred to as a turbine runner.
  • a compressor 105 for example, a toroidal combustion chamber 110, in particular annular combustion chamber, with a plurality of coaxially arranged burners 107, a turbine 108 and the exhaust housing 109th
  • the annular combustion chamber 110 communicates with an annular annular hot gas channel 111, for example.
  • Each turbine stage 112 is formed, for example, from two blade rings. As seen in the direction of flow of a working medium 113, in the hot gas channel 111 of a row of guide vanes 115, a series 125 formed of rotor blades 120 follows.
  • the guide vanes 130 are fastened to an inner housing 138 of a stator 143, whereas the moving blades 120 of a row 125 are attached to the rotor 103 by means of a turbine disk 133, for example. Coupled to the rotor 103 is a generator or work machine (not shown).
  • air 105 is sucked in and compressed by the compressor 105 through the intake housing 104.
  • the provided at the turbine end of the compressor 105 compressed air is fed to the burners 107 where it is mixed with a fuel.
  • the mixture is then burned to form the working fluid 113 in the combustion chamber 110.
  • the working medium 113 flows along the hot gas channel 111 past the guide vanes 130 and the rotor blades 120.
  • the working medium 113 expands in a pulse-transmitting manner, so that the rotor blades 120 drive the rotor 103 and drive the machine coupled to it.
  • the components exposed to the hot working medium 113 are subject to thermal loads during operation of the gas turbine 100.
  • the guide vanes 130 and rotor blades 120 of the first turbine stage 112, viewed in the flow direction of the working medium 113, are subjected to the greatest thermal stress in addition to the heat shield elements lining the annular combustion chamber 110. To withstand the prevailing temperatures, they can be cooled by means of a coolant.
  • substrates of the components can have a directional structure, ie they are monocrystalline (SX structure) or have only longitudinal grains (DS structure).
  • As the material for the components, in particular for the turbine blade 120, 130 and components of the combustion chamber 110 for example, iron-, nickel- or cobalt-based superalloys are used. Such superalloys are for example from EP 1 204 776 B1 .
  • EP 1 306 454 EP 1 319 729 A1 .
  • the blades 120, 130 may be anti-corrosion coatings (MCrAlX; M is at least one element of the group iron (Fe), cobalt (Co), nickel (Ni), X is an active element and is yttrium (Y) and / or silicon , Scandium (Sc) and / or at least one element of the rare earth or hafnium).
  • M is at least one element of the group iron (Fe), cobalt (Co), nickel (Ni)
  • X is an active element and is yttrium (Y) and / or silicon , Scandium (Sc) and / or at least one element of the rare earth or hafnium).
  • Such alloys are known from the EP 0 486 489 B1 .
  • EP 0 786 017 B1 EP 0 786 017 B1 .
  • EP 0 412 397 B1 or EP 1 306 454 A1 On the MCrAIX may still be present a thermal barrier coating, and consists for example of ZrO 2 , Y 2 O 3 -ZrO 2 , that is, it is not, partially or completely stabilized by yttria and / or calcium oxide and / or magnesium oxide.
  • suitable coating processes such as electron beam evaporation (EB-PVD), stalk-shaped grains are produced in the thermal barrier coating.
  • the vane 130 has a guide vane foot (not shown here) facing the inner housing 138 of the turbine 108 and a vane head opposite the vane foot.
  • the vane head faces the rotor 103 and fixed to a mounting ring 140 of the stator 143.
  • the FIG. 7 shows a combustion chamber 110 of a gas turbine.
  • the combustion chamber 110 is designed, for example, as a so-called annular combustion chamber, in which a plurality of burners 107 arranged around a rotation axis 102 in the circumferential direction open into a common combustion chamber space 154, which generate flames 156.
  • the combustion chamber 110 is configured in its entirety as an annular structure, which is positioned around the axis of rotation 102 around.
  • the combustion chamber 110 is designed for a comparatively high temperature of the working medium M of about 1000 ° C to 1600 ° C.
  • the combustion chamber wall 153 is provided on its side facing the working medium M side with an inner lining formed from heat shield elements 155.
  • Each heat shield element 155 made of an alloy is equipped on the working medium side with a particularly heat-resistant protective layer (MCrAIX layer and / or ceramic coating) or is made of high-temperature-resistant material (solid ceramic stones).
  • M is at least one element of the group iron (Fe), cobalt (Co), nickel (Ni), X is an active element and stands for yttrium (Y) and / or silicon and / or at least one element of the rare earths, or hafnium (Hf).
  • MCrAIX means: M is at least one element of the group iron (Fe), cobalt (Co), nickel (Ni), X is an active element and stands for yttrium (Y) and / or silicon and / or at least one element of the rare earths, or hafnium (Hf).
  • Such alloys are known from the EP 0 486 489 B1 .
  • EP 0 412 397 B1 or EP 1 306 454 A1 are known from the EP 0 486 489 B1 .
  • EP 0 412 397 B1 or EP 1 306 454 A1 is known from the EP 0 486 489 B1 .
  • a ceramic thermal barrier coating may be present and consists for example of ZrO 2 , Y 2 O 3 -ZrO 2 , ie it is not, partially or completely stabilized by yttria and / or calcium oxide and / or magnesium oxide.
  • suitable coating processes such as electron beam evaporation (EB-PVD)
  • stalk-shaped grains are produced in the thermal barrier coating.
  • APS atmospheric plasma spraying
  • LPPS LPPS
  • VPS vacuum plasma spraying
  • CVD chemical vaporation
  • the thermal barrier coating may have porous, micro- or macro-cracked grains for better thermal shock resistance.
  • Refurbishment means that heat shield elements 155 may need to be deprotected (e.g., by sandblasting) after use. This is followed by removal of the corrosion and / or oxidation layers or products. If necessary, cracks in the heat shield element 155 are also repaired. This is followed by a recoating of the heat shield elements 155 and a renewed use of the heat shield elements 155.
  • the heat shield elements 155 are then, for example, hollow and possibly still have cooling holes (not shown) which open into the combustion chamber space 154.
  • FIG. 8 shows by way of example a gas turbine 100 in a longitudinal partial section.
  • the gas turbine 100 has inside a rotatably mounted about a rotation axis 102 rotor 103 with a shaft 101, which is also referred to as a turbine runner.
  • a compressor 105 for example, a toroidal combustion chamber 110, in particular annular combustion chamber, with a plurality of coaxially arranged burners 107, a turbine 108 and the exhaust housing 109th
  • the annular combustion chamber 110 communicates with an annular annular hot gas channel 111, for example.
  • Each turbine stage 112 is formed, for example, from two blade rings. As seen in the direction of flow of a working medium 113, in the hot gas channel 111 of a row of guide vanes 115, a series 125 formed of rotor blades 120 follows.
  • the guide vanes 130 are fastened to an inner housing 138 of a stator 143, whereas the moving blades 120 of a row 125 are attached to the rotor 103 by means of a turbine disk 133, for example. Coupled to the rotor 103 is a generator or work machine (not shown).
  • air 105 is sucked in and compressed by the compressor 105 through the intake housing 104.
  • the compressed air provided at the turbine-side end of the compressor 105 is supplied to the burners 107 where it is mixed with a fuel.
  • the mixture is then burned to form the working fluid 113 in the combustion chamber 110.
  • the working medium 113 flows along the hot gas channel 111 past the guide vanes 130 and the rotor blades 120. Relaxed on the rotor blades 120 the working medium 113 is pulse-transmitting, so that the rotor blades 120 drive the rotor 103 and this drives the machine coupled to it.
  • the components exposed to the hot working medium 113 are subject to thermal loads during operation of the gas turbine 100.
  • the guide vanes 130 and rotor blades 120 of the first turbine stage 112, viewed in the flow direction of the working medium 113, are subjected to the greatest thermal stress in addition to the heat shield elements lining the annular combustion chamber 110. To withstand the prevailing temperatures, they can be cooled by means of a coolant.
  • substrates of the components can have a directional structure, ie they are monocrystalline (SX structure) or have only longitudinal grains (DS structure).
  • As the material for the components, in particular for the turbine blade 120, 130 and components of the combustion chamber 110 for example, iron-, nickel- or cobalt-based superalloys are used. Such superalloys are for example from EP 1 204 776 B1 .
  • EP 1 306 454 EP 1 319 729 A1 .
  • the blades 120, 130 may have anticorrosive coatings (MCrAIX; M is at least one element of the group iron (Fe), cobalt (Co), nickel (Ni), X is an active element and represents yttrium (Y) and / or silicon , Scandium (Sc) and / or at least one element of the rare earth or hafnium).
  • MCrAIX anticorrosive coatings
  • Such alloys are known from the EP 0 486 489 B1 .
  • EP 0 412 397 B1 or EP 1 306 454 A1 are known from the EP 0 486 489 B1 .
  • EP 0 786 017 B1 EP 0 412 397 B1 or EP 1 306 454 A1 .
  • MCrAlX may still be present a thermal barrier coating, and consists for example of ZrO 2 , Y 2 O 3 -ZrO 2 , that is, it is not, partially or completely stabilized by yttria and / or calcium oxide and / or magnesium oxide.
  • Electron beam evaporation produces stalk-shaped grains in the thermal barrier coating.
  • the vane 130 has a guide vane foot (not shown here) facing the inner housing 138 of the turbine 108 and a vane head opposite the vane foot.
  • the vane head faces the rotor 103 and fixed to a mounting ring 140 of the stator 143.

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  • Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Mechanical Engineering (AREA)
  • Materials Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Chemical Kinetics & Catalysis (AREA)
  • Metallurgy (AREA)
  • Organic Chemistry (AREA)
  • Ceramic Engineering (AREA)
  • Combustion & Propulsion (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
  • Other Surface Treatments For Metallic Materials (AREA)
EP11188032.4A 2011-11-07 2011-11-07 Couche d'isolation thermique en céramique sur une surface structurée et procédé de fabrication Withdrawn EP2589682A1 (fr)

Priority Applications (4)

Application Number Priority Date Filing Date Title
EP11188032.4A EP2589682A1 (fr) 2011-11-07 2011-11-07 Couche d'isolation thermique en céramique sur une surface structurée et procédé de fabrication
US14/354,573 US9862002B2 (en) 2011-11-07 2012-09-14 Process for producing a layer system
EP12759691.4A EP2753729A1 (fr) 2011-11-07 2012-09-14 Procédé de fabrication d'un système en couches
PCT/EP2012/068048 WO2013068159A1 (fr) 2011-11-07 2012-09-14 Procédé de fabrication d'un système en couches

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
EP11188032.4A EP2589682A1 (fr) 2011-11-07 2011-11-07 Couche d'isolation thermique en céramique sur une surface structurée et procédé de fabrication

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EP2589682A1 true EP2589682A1 (fr) 2013-05-08

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EP11188032.4A Withdrawn EP2589682A1 (fr) 2011-11-07 2011-11-07 Couche d'isolation thermique en céramique sur une surface structurée et procédé de fabrication
EP12759691.4A Withdrawn EP2753729A1 (fr) 2011-11-07 2012-09-14 Procédé de fabrication d'un système en couches

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EP12759691.4A Withdrawn EP2753729A1 (fr) 2011-11-07 2012-09-14 Procédé de fabrication d'un système en couches

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US (1) US9862002B2 (fr)
EP (2) EP2589682A1 (fr)
WO (1) WO2013068159A1 (fr)

Cited By (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
DE102015222808A1 (de) * 2015-11-19 2017-05-24 Siemens Aktiengesellschaft Segmentiertes zweilagiges Schichtsystem
DE102015222812A1 (de) * 2015-11-19 2017-05-24 Siemens Aktiengesellschaft Keramisches Schichtsystem mit Vertiefungen in keramischer Schicht und strukturierter Haftvermittlerschicht
EP3222747A1 (fr) * 2016-03-24 2017-09-27 Siemens Aktiengesellschaft Composant de gaz chaud

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* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP2733310A1 (fr) * 2012-11-16 2014-05-21 Siemens Aktiengesellschaft Surface modifiée autour d'un trou
JP6065163B1 (ja) * 2015-03-18 2017-01-25 中国電力株式会社 高温部品のひずみ測定方法及び高温部品
DE102015224844A1 (de) * 2015-12-10 2017-06-14 Siemens Aktiengesellschaft Bauteil mit lokaler Verstärkung bezüglich Festigkeit und Oxidationsbeständigkeit und Verfahren

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Publication number Priority date Publication date Assignee Title
EP0486489B1 (fr) 1989-08-10 1994-11-02 Siemens Aktiengesellschaft Revetement anticorrosion resistant aux temperatures elevees, notamment pour elements de turbines a gaz
US5419971A (en) * 1993-03-03 1995-05-30 General Electric Company Enhanced thermal barrier coating system
EP0412397B1 (fr) 1989-08-10 1998-03-25 Siemens Aktiengesellschaft Revêtement protecteur contenant du rhénium possédant une résistance plus grande à la corrosion et l'oxydation
EP0786017B1 (fr) 1994-10-14 1999-03-24 Siemens Aktiengesellschaft Couche de protection de pieces contre la corrosion, l'oxydation et les contraintes thermiques excessives, et son procede de production
WO1999067435A1 (fr) 1998-06-23 1999-12-29 Siemens Aktiengesellschaft Alliage a solidification directionnelle a resistance transversale a la rupture amelioree
US6074706A (en) * 1998-12-15 2000-06-13 General Electric Company Adhesion of a ceramic layer deposited on an article by casting features in the article surface
WO2000044949A1 (fr) 1999-01-28 2000-08-03 Siemens Aktiengesellschaft Superalliage a base de nickel presentant une bonne usinabilite
EP1306454A1 (fr) 2001-10-24 2003-05-02 Siemens Aktiengesellschaft Revêtement protecteur contenant du rhénium pour la protection d'un élément contre l'oxydation et la corrosion aux températures élevées
EP1319729A1 (fr) 2001-12-13 2003-06-18 Siemens Aktiengesellschaft Pièce résistante à des températures élevées réalisé en superalliage polycristallin ou monocristallin à base de nickel
WO2004043691A1 (fr) * 2002-11-12 2004-05-27 University Of Virginia Patent Foundation Revetement de protection thermique extremement resistant aux contraintes, et procede et dispositif associes
EP1204776B1 (fr) 1999-07-29 2004-06-02 Siemens Aktiengesellschaft Piece resistant a des temperatures elevees et son procede de production
US20080085191A1 (en) * 2006-10-05 2008-04-10 Siemens Power Generation, Inc. Thermal barrier coating system for a turbine airfoil usable in a turbine engine
US20090017260A1 (en) * 2001-08-02 2009-01-15 Kulkarni Anand A Segmented thermal barrier coating
EP2275645A2 (fr) * 2009-07-17 2011-01-19 Rolls-Royce Corporation Composant de turbine à gaz comprenant des caractéristiques de réduction de la fatigue

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US6846574B2 (en) * 2001-05-16 2005-01-25 Siemens Westinghouse Power Corporation Honeycomb structure thermal barrier coating

Patent Citations (14)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP0486489B1 (fr) 1989-08-10 1994-11-02 Siemens Aktiengesellschaft Revetement anticorrosion resistant aux temperatures elevees, notamment pour elements de turbines a gaz
EP0412397B1 (fr) 1989-08-10 1998-03-25 Siemens Aktiengesellschaft Revêtement protecteur contenant du rhénium possédant une résistance plus grande à la corrosion et l'oxydation
US5419971A (en) * 1993-03-03 1995-05-30 General Electric Company Enhanced thermal barrier coating system
EP0786017B1 (fr) 1994-10-14 1999-03-24 Siemens Aktiengesellschaft Couche de protection de pieces contre la corrosion, l'oxydation et les contraintes thermiques excessives, et son procede de production
WO1999067435A1 (fr) 1998-06-23 1999-12-29 Siemens Aktiengesellschaft Alliage a solidification directionnelle a resistance transversale a la rupture amelioree
US6074706A (en) * 1998-12-15 2000-06-13 General Electric Company Adhesion of a ceramic layer deposited on an article by casting features in the article surface
WO2000044949A1 (fr) 1999-01-28 2000-08-03 Siemens Aktiengesellschaft Superalliage a base de nickel presentant une bonne usinabilite
EP1204776B1 (fr) 1999-07-29 2004-06-02 Siemens Aktiengesellschaft Piece resistant a des temperatures elevees et son procede de production
US20090017260A1 (en) * 2001-08-02 2009-01-15 Kulkarni Anand A Segmented thermal barrier coating
EP1306454A1 (fr) 2001-10-24 2003-05-02 Siemens Aktiengesellschaft Revêtement protecteur contenant du rhénium pour la protection d'un élément contre l'oxydation et la corrosion aux températures élevées
EP1319729A1 (fr) 2001-12-13 2003-06-18 Siemens Aktiengesellschaft Pièce résistante à des températures élevées réalisé en superalliage polycristallin ou monocristallin à base de nickel
WO2004043691A1 (fr) * 2002-11-12 2004-05-27 University Of Virginia Patent Foundation Revetement de protection thermique extremement resistant aux contraintes, et procede et dispositif associes
US20080085191A1 (en) * 2006-10-05 2008-04-10 Siemens Power Generation, Inc. Thermal barrier coating system for a turbine airfoil usable in a turbine engine
EP2275645A2 (fr) * 2009-07-17 2011-01-19 Rolls-Royce Corporation Composant de turbine à gaz comprenant des caractéristiques de réduction de la fatigue

Cited By (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
DE102015222808A1 (de) * 2015-11-19 2017-05-24 Siemens Aktiengesellschaft Segmentiertes zweilagiges Schichtsystem
DE102015222812A1 (de) * 2015-11-19 2017-05-24 Siemens Aktiengesellschaft Keramisches Schichtsystem mit Vertiefungen in keramischer Schicht und strukturierter Haftvermittlerschicht
EP3222747A1 (fr) * 2016-03-24 2017-09-27 Siemens Aktiengesellschaft Composant de gaz chaud

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EP2753729A1 (fr) 2014-07-16
WO2013068159A1 (fr) 2013-05-16
US9862002B2 (en) 2018-01-09
US20140295086A1 (en) 2014-10-02

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