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WO2014108199A1 - Procédé de fabrication de turbines à gaz et procédé permettant de faire fonctionner un système de turbines à gaz - Google Patents

Procédé de fabrication de turbines à gaz et procédé permettant de faire fonctionner un système de turbines à gaz Download PDF

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Publication number
WO2014108199A1
WO2014108199A1 PCT/EP2013/050452 EP2013050452W WO2014108199A1 WO 2014108199 A1 WO2014108199 A1 WO 2014108199A1 EP 2013050452 W EP2013050452 W EP 2013050452W WO 2014108199 A1 WO2014108199 A1 WO 2014108199A1
Authority
WO
WIPO (PCT)
Prior art keywords
layer
ceramic
thermal insulation
new
metallic protective
Prior art date
Application number
PCT/EP2013/050452
Other languages
German (de)
English (en)
Inventor
Axel Kaiser
Werner Stamm
Original Assignee
Siemens Aktiengesellschaft
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 Aktiengesellschaft filed Critical Siemens Aktiengesellschaft
Priority to PCT/EP2013/050452 priority Critical patent/WO2014108199A1/fr
Priority to EP13700374.5A priority patent/EP2904129A1/fr
Publication of WO2014108199A1 publication Critical patent/WO2014108199A1/fr

Links

Classifications

    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/12—Blades
    • F01D5/28—Selecting particular materials; Particular measures relating thereto; Measures against erosion or corrosion
    • F01D5/288—Protective coatings for blades
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B23—MACHINE TOOLS; METAL-WORKING NOT OTHERWISE PROVIDED FOR
    • B23K—SOLDERING OR UNSOLDERING; WELDING; CLADDING OR PLATING BY SOLDERING OR WELDING; CUTTING BY APPLYING HEAT LOCALLY, e.g. FLAME CUTTING; WORKING BY LASER BEAM
    • B23K10/00—Welding or cutting by means of a plasma
    • B23K10/02—Plasma welding
    • B23K10/027—Welding for purposes other than joining, e.g. build-up welding
    • 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
    • C23C28/00—Coating for obtaining at least two superposed coatings either by methods not provided for in a single one of groups C23C2/00 - C23C26/00 or by combinations of methods provided for in subclasses C23C and C25C or C25D
    • 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
    • C23C28/00—Coating for obtaining at least two superposed coatings either by methods not provided for in a single one of groups C23C2/00 - C23C26/00 or by combinations of methods provided for in subclasses C23C and C25C or C25D
    • C23C28/30—Coatings combining at least one metallic layer and at least one inorganic non-metallic layer
    • 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
    • B23K—SOLDERING OR UNSOLDERING; WELDING; CLADDING OR PLATING BY SOLDERING OR WELDING; CUTTING BY APPLYING HEAT LOCALLY, e.g. FLAME CUTTING; WORKING BY LASER BEAM
    • B23K2101/00—Articles made by soldering, welding or cutting
    • B23K2101/001—Turbines
    • 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
    • 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
    • F05D2300/00—Materials; Properties thereof
    • F05D2300/20—Oxide or non-oxide ceramics
    • 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
    • F05D2300/00—Materials; Properties thereof
    • F05D2300/50—Intrinsic material properties or characteristics
    • F05D2300/514—Porosity

Definitions

  • the invention relates to a method for the production of gas turbines, which are designed to be flexible and methods for operating gas turbines.
  • Gas turbines can be operated during power generation in base load operation or in particular in peak load operation.
  • the object is achieved by a method for producing gas turbines according to claim 1 and a method according to claim 33.
  • the metallic adhesion promoter ⁇ layer and the ceramic layer are changed to it
  • a two-layer ceramic thermal barrier coating (13) is removed from the first turbine blades and / or
  • a single-layer thermal barrier coating ⁇ 1 '' is applied as a second ceramic thermal barrier coating on the second or new second turbine blades - in which the single-layer ceramic thermal barrier coating ⁇ 1 '') is produced with a porosity of 18% ⁇ 4%,
  • a thinner ceramic thermal barrier coating (7) replaces the first ceramic thermal barrier layer with a thicker ceramic thermal barrier coating (7 ', 13') as a second ceramic thermal barrier coating of the second or new two ⁇ th turbine blades
  • a thicker ceramic thermal barrier coating as the first Kerami ⁇ specific heat-insulating layer is replaced by a thinner Kerami ⁇ specific thermal insulation layer as a second ceramic thermal barrier layer of the second or the new second turbine blades, wherein the difference in thickness is at least -50 ⁇
  • the two-layer thermal barrier coating (13 ') with a lowermost ceramic layer ⁇ !' '') is produced with a porosity of 12% ⁇ 4% and with an outer ceramic layer (10 ') with a porosity of 18% ⁇ 4%,
  • the lower layer ⁇ ! '' ') of the two-layer thermal barrier coating (13) is made thinner, in particular at least 20% thinner than the upper layer (10'),
  • the lower layer ⁇ ! ''' has a thickness of 75ym the two-layer thermal barrier coating (13) to 150ym, more in particular, the total thickness of the two-layer Wär ⁇ medämm harsh (13) is 500ym to 800ym
  • Zirconium oxide is used for the ceramic thermal barrier coating (7 ', 7' ', 13') or the ceramic layers ⁇ ! '' ', 10'), and the monoclinic fraction of the powder to be sprayed under o,
  • the lower layer ⁇ 1 '' ') is sprayed without polymer and the upper layer (10') is sprayed with polymer
  • the average pore diameter (Dio) of the upper ceramic layer (10 ') is generated larger as the average pore diameter (d 7 ) of the lower ceramic layer ⁇ ! '''),
  • Ni nickel
  • the single-layer metallic protective layer contains the following elements
  • the single-layer metallic protective layer having the following composition (in% by weight):
  • Chromium content of the substrate and the chromium content of the outer metallic protective layer Chromium content of the substrate and the chromium content of the outer metallic protective layer
  • the metallic protective layers comprise an alloy MCrAlX
  • the outer protective layer comprises tantalum (Ta) or tantalum (Ta) and iron (Fe),
  • the lower NiCoCrAlY layer has the following composition ⁇ (% in wt .-):
  • the upper NiCoCrAlY layer has the following composition (in% by weight):
  • Figure 4 shows a pore distribution in a ceramic
  • FIG. 5 shows 8 exemplary embodiments of the invention
  • Figure 10 is a gas turbine. The description is only an exemplary embodiment of the invention ⁇ .
  • An interval of maintenance of gas turbines 100 ( Figure 6) is determined by detecting operating hours and starts, which are dependent on the mode of operation and certain factors. Maintenance must be carried out after reaching the hourly or start limit. If maintenance is required depending on the field of application of the gas turbine, or if the use requires an overhaul or another use beforehand, then the configuration of the gas turbine 100 is changed. Definition of terms:
  • First gas turbine has 1st turbine blade with 1st metallic protective layer.
  • Second gas turbine has turbine blades with metallic protective coating
  • First gas turbine has 1st turbine blade with 1st thermal insulation ⁇ layer.
  • Second gas turbine has turbine blades with ceramic thermal barrier coatings
  • a single-layer metallic protective layer, especially with the composition NiCoCrAlX is suitable for a "Daily Starter".
  • the change in the layer always means a change in the chemical composition, especially because a lower and upper metallic protective layer differ significantly in their composition, ie. at least one alloying element has an at least 15% Various ⁇ NEN portion.
  • Figure 5 shows a substrate 4 having a single-layer metalli ⁇ rule protective layer 16 is removed, and is changed for the production of second turbine acting ⁇ feln or new second turbine blades in its layer thickness 16 'of a first turbine blade.
  • the alloy of the original metallic protective layer 16 may be altered.
  • Figure 6 it is shown that a two-layer metallic protective layer 22 ', 22''from the first turbine blades is removed and a single-layer metal as a new second protective layer is applied ⁇ 22nd
  • the layer thickness of the single-layer metallic protective layer 22 may preferably correspond to that of the two metallic protective layers 22 ', 22 ". At least one alloy of one of the two metallic protective layers 22 ', 22 "may also correspond to the alloy of the single-layer layer 22 or may be entirely different from one another.
  • the layer thickness can change.
  • FIG. 8 shows that a single-layer metallic protective layer 16 is removed from a substrate 4 of first turbine blades and a two-layer metallic new protective second layer 24 ', 24 "is applied.
  • the layer thickness of the two-layer metallic protective layer 24 ', 24 " may preferably correspond to that of the original single-layer metallic protective layer 16.
  • the two alloys are the two-ply
  • the variation of the metallic protective layers according to FIGS. 5, 6, 7 and / or 8 can be varied with the variation of the ceramic protective layers according to FIGS. 1, 2 and 3.
  • a preferred metallic single-layer protective layer comprises:
  • composition having the following composition (in% by weight):
  • the single layer metallic protective layer comprises: 29% - 31% nickel (Ni),
  • a two-layer MCrAlX layer 24 ', 24'' is suitable for long-term operation (base loader).
  • a two-layer metallic protective layer 24 ', 24 "based on the NiCoCrAlX alloys it is preferable to use a tantalum or tantalum / iron-containing outer metallic protective layer.
  • An advantageous composition for the outer protective layer results from (in% by weight)
  • Ni NiCoCrAlX layer
  • the lower NiCoCrAlY layer has the following composition (in% by weight):
  • the turbine blades for the second gas turbine may be at the origin (same substrate) the first turbine blades of the first gas turbine or other gas turbines already in use, refurbished accordingly, and re-coating second turbine blades
  • the gas turbine 100 has a two-layer ceramic thermal barrier coating on the turbine blades 120, 130, to apply a single-layer TBC so that it can then be used in peak load mode (FIG. 2).
  • a ceramic layer is used with a uniform porosity.
  • the ceramic thermal barrier coating on the turbine blades 120, 130 preferential ⁇ as a high porosity of 18% ⁇ 4%.
  • Each ceramic sprayed layer is applied in coating layers.
  • Two-ply means, however, that a second layer differs from a first, lower layer due to porosity and / or microstructure and / or chemical composition.
  • the lower layer used is preferably a ceramic layer 7 with a porosity of 12% ⁇ 4%, which preferably has a layer thickness of 75 ⁇ m to 150 ⁇ m.
  • a porosity of 18% ⁇ 4% is sprayed or is present as outer ceramic layer 10.
  • the difference in porosity is at least 2%, especially at least 4%. Variations in the porosity in the production are known. Within a batch, i. of a blade set, there are no fluctuations.
  • a ceramic layer is also preferably 7 having a porosity of 12% ⁇ 4%, the preference ⁇ a layer thickness of 75ym having up 150ym.
  • a porosity of 18% ⁇ 4% is sprayed or is present as outer ceramic layer 10.
  • a ceramic layer is also preferably 7 having a porosity of 18% ⁇ 4%, the preference ⁇ a layer thickness of up to 75ym having 150ym.
  • coarse grains may be used in the spraying and polymer use or smaller grains may be used, wherein roughly means at least 20% larger average particle diameter.
  • a two-layer ceramic layer 7, 10 can be manufactured with Various ⁇ NEN Spraying: the lower layer 7 is without polymer and the upper layer 10 is injected ⁇ ver with polymer.
  • the same powder is used, so also a same particle size distribution.
  • Zirconium oxide (ZrÜ 2) for the ceramic layers of the thermal barrier ⁇ layers preferably has a monoclinic fraction of ⁇ 3%, in particular -S 1.5%. Corresponding portions then has a ceramic layer or layer 7, 7 ', 10, 13 (FIGS. 1-3) on the turbine blade 120, 130.
  • the minimum proportion of the monoclinic phase is at least 1%, in particular 0.5%, in order not to increase the cost of the powder too much.
  • FIG. 9 shows a perspective view of a rotor blade or guide vane 130 of a turbomachine that extends along a longitudinal axis 121.
  • the turbomachine may be a gas turbine of an aircraft or a power plant for electricity generation, a steam turbine or a compressor.
  • the blade 120, 130 has along the longitudinal axis 121 to each other, a securing region 400, an adjoining blade or vane platform 403 and a blade 406 and a blade tip 415.
  • the vane 130 As a guide vane 130, the vane 130 having at its blade tip 415 have a further platform (not Darge ⁇ asserted).
  • a blade root 183 is formed, which serves for attachment of the blades 120, 130 to a shaft or a disc (not shown).
  • the blade root 183 is, for example, as a hammerhead out staltet ⁇ . Other designs as fir tree or Schissebwschwanzfuß are possible.
  • the blade 120, 130 has for a medium which flows past the scene ⁇ felblatt 406 on a leading edge 409 and a trailing edge 412th
  • conventional blades 120, 130 in all regions 400, 403, 406 of the blade 120, 130, for example, massive metallic materials, in particular superalloys, are used.
  • Such superalloys are known, for example, from EP 1 204 776 B1, EP 1 306 454, EP 1 319 729 A1, WO 99/67435 or WO 00/44949.
  • the blade 120, 130 can be made by a casting process, also by directional solidification, by a forging process, by a milling process or combinations thereof.
  • Workpieces with a monocrystalline structure or structures are used as components for machines which are exposed to high mechanical, thermal and / or chemical stresses during operation.
  • Such monocrystalline workpieces takes place e.g. by directed solidification from the melt.
  • These are casting processes in which the liquid metallic alloy is transformed into a monocrystalline structure, i. to the single-crystal workpiece, or directionally solidified.
  • dendritic crystals are aligned along the heat flow and form either a columnar grain structure (columnar, ie grains that run the entire length of the workpiece and here, in common parlance, referred to as directionally solidified) or a monocrystalline structure, ie the whole workpiece be ⁇ is made of a single crystal.
  • a columnar grain structure columnar, ie grains that run the entire length of the workpiece and here, in common parlance, referred to as directionally solidified
  • a monocrystalline structure ie the whole workpiece be ⁇ is made of a single crystal.
  • directionally solidified microstructures refers both to single crystals which have no grain boundaries or at most small-angle grain boundaries, and to stem crystal structures which are likely to be found in longitudinal grooves. grain boundaries, but have no transverse grain boundaries. These second-mentioned crystalline structures are also known as directionally solidified structures.
  • the blades 120, 130 may have coatings against corrosion or oxidation, e.g. B. (MCrAlX; M is at least one element of the group iron (Fe), cobalt (Co),
  • Nickel (Ni) 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 EP 0 486 489 B1, EP 0 786 017 B1, EP 0 412 397 B1 or EP 1 306 454 A1.
  • the density is preferably 95% of the theoretical
  • the layer composition comprises Co-30Ni-28Cr-8A1-0, 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-IIAl-O, 4Y-2Re or Ni-25Co-17Cr-10A1-0, 4Y-1 are also preferably used , 5Re.
  • thermal barrier coating which is preferably the outermost layer, and consists for example of Zr0 2 , Y2Ü3-Zr02, ie it is not, partially ⁇ or fully stabilized by yttria
  • the thermal barrier coating covers the entire MCrAlX layer.
  • Suitable coating processes such as electron beam evaporation (EB-PVD), produce stalk-shaped grains in the thermal barrier coating.
  • Other coating methods are conceivable, for example atmospheric plasma spraying (APS), LPPS, VPS or CVD.
  • APS atmospheric plasma spraying
  • LPPS LPPS
  • VPS VPS
  • CVD chemical vapor deposition
  • the heat insulation layer may have ⁇ porous, micro- or macro-cracked compatible grains for better thermal shock resistance.
  • the thermal barrier coating is therefore preferably more porous than the
  • Refurbishment means that components 120, 130 may have to be freed from protective layers after use (eg by sandblasting). This is followed by removal of the corrosion and / or oxidation layers or products. Optionally, even cracks in the component 120, 130 are repaired. Thereafter, a ⁇ As the coating of the component 120, 130, after entry set of the component 120, the 130th
  • the blade 120, 130 may be hollow or solid. If the blade 120, 130 is to be cooled, it is hollow and may still have film cooling holes 418 (indicated by dashed lines).
  • FIG. 10 shows by way of example a gas turbine 100 in a longitudinal partial section.
  • the gas turbine 100 has a rotatably mounted about a rotational axis 102 ⁇ rotor 103 having a shaft 101, which is also referred to as the turbine rotor.
  • an intake housing 104 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
  • 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.
  • annular annular hot gas channel 111 for example.
  • turbine stages 112 connected in series form the turbine 108.
  • Each turbine stage 112 is formed, for example, from two blade rings .
  • a row 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.
  • air 135 is sucked by the compressor 105 through the intake housing and ver ⁇ seals.
  • the 105 ⁇ be compressed air provided at the turbine end of the compressor 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.
  • the working medium 113 expands on the rotor blades 120 in a pulse-transmitting manner, so that the rotor blades 120 drive the rotor 103 and drive the machine connected 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 highest thermal stress in addition to the heat shield elements lining the annular combustion chamber 110.
  • substrates of the components may have a directional structure, i. they are monocrystalline (SX structure) or have only longitudinal grains (DS structure).
  • the components in particular for the turbine ⁇ blade 120, 130 and components of the combustion chamber 110 are For example, iron-, nickel- or cobalt-based superalloys used.
  • Such superalloys are known, for example, from EP 1 204 776 B1, EP 1 306 454, EP 1 319 729 A1, WO 99/67435 or WO 00/44949.
  • 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 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 Zr02, Y203-Zr02, ie it is not, partially or completely stabilized by Ytt ⁇ riumoxid and / or calcium oxide and / or magnesium oxide.
  • Electron beam evaporation produces stalk-shaped grains in the thermal barrier coating.
  • the guide blade 130 has a guide blade root facing the inner housing 138 of the turbine 108 (not shown here) and a guide blade foot opposite
  • 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)
  • Chemical Kinetics & Catalysis (AREA)
  • Metallurgy (AREA)
  • Organic Chemistry (AREA)
  • General Engineering & Computer Science (AREA)
  • Inorganic Chemistry (AREA)
  • Physics & Mathematics (AREA)
  • Plasma & Fusion (AREA)
  • Coating By Spraying Or Casting (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)

Abstract

L'invention concerne un procédé de fabrication de turbines à gaz et procédé permettant de faire fonctionner un système de turbines à gaz. L'utilisation de différentes couches de céramique permet de produire différentes configurations de turbines à gaz qui sont ensuite optimisées pour un domaine d'utilisation respectivement en mode en charge de base ou en mode en pointe de charge.
PCT/EP2013/050452 2013-01-11 2013-01-11 Procédé de fabrication de turbines à gaz et procédé permettant de faire fonctionner un système de turbines à gaz WO2014108199A1 (fr)

Priority Applications (2)

Application Number Priority Date Filing Date Title
PCT/EP2013/050452 WO2014108199A1 (fr) 2013-01-11 2013-01-11 Procédé de fabrication de turbines à gaz et procédé permettant de faire fonctionner un système de turbines à gaz
EP13700374.5A EP2904129A1 (fr) 2013-01-11 2013-01-11 Procédé de fabrication de turbines à gaz et procédé permettant de faire fonctionner un système de turbines à gaz

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
PCT/EP2013/050452 WO2014108199A1 (fr) 2013-01-11 2013-01-11 Procédé de fabrication de turbines à gaz et procédé permettant de faire fonctionner un système de turbines à gaz

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Publication Number Publication Date
WO2014108199A1 true WO2014108199A1 (fr) 2014-07-17

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WO (1) WO2014108199A1 (fr)

Cited By (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
WO2017089082A1 (fr) * 2015-11-27 2017-06-01 Siemens Aktiengesellschaft Barrière thermique double couche locale

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JP2019504233A (ja) * 2015-11-27 2019-02-14 シーメンス アクティエンゲゼルシャフト 局所２層遮熱コーティング
US10662787B2 (en) 2015-11-27 2020-05-26 Siemens Aktiengesellschaft Local two-layer thermal barrier coating

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