WO2022035460A1 - Revêtement formant barrière thermique, élément stratifié comprenant un revêtement formant barrière thermique et turbine à gaz comprenant un élément stratifié - Google Patents

Revêtement formant barrière thermique, élément stratifié comprenant un revêtement formant barrière thermique et turbine à gaz comprenant un élément stratifié Download PDF

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Publication number
WO2022035460A1
WO2022035460A1 PCT/US2020/070411 US2020070411W WO2022035460A1 WO 2022035460 A1 WO2022035460 A1 WO 2022035460A1 US 2020070411 W US2020070411 W US 2020070411W WO 2022035460 A1 WO2022035460 A1 WO 2022035460A1
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WO
WIPO (PCT)
Prior art keywords
barrier coating
thermal barrier
density
mean
turbine engine
Prior art date
Application number
PCT/US2020/070411
Other languages
English (en)
Inventor
Atin SHARMA
Ramesh Subramanian
Rohit KHATTAR
Original Assignee
Siemens Gas And Power Gmbh & Co. Kg
Siemens Energy, Inc.
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 Gas And Power Gmbh & Co. Kg, Siemens Energy, Inc. filed Critical Siemens Gas And Power Gmbh & Co. Kg
Priority to PCT/US2020/070411 priority Critical patent/WO2022035460A1/fr
Publication of WO2022035460A1 publication Critical patent/WO2022035460A1/fr

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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
    • F01D5/12—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
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D9/00—Stators
    • F01D9/02—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles
    • F01D9/023—Transition ducts between combustor cans and first stage of the turbine in gas-turbine engines; their cooling or sealings
    • 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
    • F05D2230/10—Manufacture by removing material
    • F05D2230/13—Manufacture by removing material using lasers
    • 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
    • F05D2240/10—Stators
    • F05D2240/11—Shroud seal segments
    • 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
    • F05D2240/10—Stators
    • F05D2240/12—Fluid guiding means, e.g. vanes
    • 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
    • F05D2240/20—Rotors
    • F05D2240/30—Characteristics of rotor blades, i.e. of any element transforming dynamic fluid energy to or from rotational energy and being attached to a rotor
    • 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
    • F05D2240/35—Combustors or associated equipment
    • 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
    • F05D2240/55—Seals
    • 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
    • F05D2240/60—Shafts
    • 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
    • F05D2250/10—Two-dimensional
    • F05D2250/11—Two-dimensional triangular
    • 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
    • F05D2250/10—Two-dimensional
    • F05D2250/12—Two-dimensional rectangular
    • 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
    • F05D2250/10—Two-dimensional
    • F05D2250/12—Two-dimensional rectangular
    • F05D2250/121—Two-dimensional rectangular square
    • 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
    • F05D2250/10—Two-dimensional
    • F05D2250/13—Two-dimensional trapezoidal
    • F05D2250/132—Two-dimensional trapezoidal hexagonal
    • 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
    • F05D2250/21—Three-dimensional pyramidal
    • 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
    • F05D2250/23—Three-dimensional prismatic
    • F05D2250/232—Three-dimensional prismatic conical
    • 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
    • F05D2250/28—Three-dimensional patterned
    • F05D2250/282—Three-dimensional patterned cubic pattern
    • 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
    • F05D2250/28—Three-dimensional patterned
    • F05D2250/283—Three-dimensional patterned honeycomb
    • 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
    • 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
    • F05D2250/60—Structure; Surface texture
    • F05D2250/62—Structure; Surface texture smooth or fine
    • 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
    • F05D2260/00—Function
    • F05D2260/94—Functionality given by mechanical stress related aspects such as low cycle fatigue [LCF] of high cycle fatigue [HCF]
    • F05D2260/941—Functionality given by mechanical stress related aspects such as low cycle fatigue [LCF] of high cycle fatigue [HCF] particularly aimed at mechanical or thermal stress reduction
    • 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
    • F05D2300/10—Metals, alloys or intermetallic compounds
    • F05D2300/17—Alloys
    • F05D2300/175—Superalloys
    • 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
    • F05D2300/21—Oxide ceramics
    • F05D2300/2118—Zirconium oxides
    • 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

  • a gas turbine engine typically includes a compressor section, a combustor section and a turbine section.
  • the compressor section typically includes multiple stages of rotating compressor blades and stationary compressor vanes that compress inflowing atmospheric air, and is operatively connected to the combustor section.
  • the combustor section typically includes a diffusor that diffuses the compressed air into a combustor which mixes the compressed air with fuel, ignites the mixture, and transits the ignited hot gas mixture along a hot gas path to the operatively connected turbine section.
  • the turbine section extracts energy from the ignited hot gas within the hot gas path via multiple stages of rotating turbine blades and stationary turbine vanes in order to generate mechanical power in industrial applications or propulsion in aero application or for other useful purposes.
  • Certain internal components of the gas turbine engine are routinely exposed to temperatures up to approximately 1250 degrees Celsius, with next generation engines likely having hot gas path temperatures up to 1400 degrees Celsius and higher.
  • these engine components exposed to the hot gas path are typically constructed of high temperature resistant superalloys.
  • These superalloy components may often include internal cooling passages for passage of coolant fluid to cool the surfaces exposed to the hot gas path.
  • These superalloy components may also often include an even higher temperature resistant ceramic thermal barrier coating applied to the surface of the superalloy to further assist the component in withstanding exposure to the high temperature environment.
  • a gas turbine engine includes a compressor section; a combustor section operatively connected to the compressor section; a turbine section operatively connected to the combustor section; a hot gas path extending from the combustor section to the turbine section; and a layered component designed and constructed for prolonged exposure to the hot gas path including a substrate and a thermal barrier coating having: a single layer of 6- 15% wt.% yttria-stabilized zirconia in a tetragonal phase arranged on or above the substrate, where the single layer does not have an adjacent upper or lower layer of a second thermal barrier coating, a density having a homogeneous distribution such that one standard deviation of the mean or median density is not more than +/- 25% of the mean or median density and having a mean bulk value density of greater than 70%, a smooth exposed surface having an Ra of less than 5 microns, and a plurality of grooves extending from the single layer surface toward the substrate with dimensions of 30 - 300 z/m
  • a layered component is designed and constructed for prolonged exposure to a hot gas path comprising: a superalloy substrate, a bond coat arranged on the superalloy substrate, and a thermal barrier coating comprising: a single layer of 6-10% wt.% yttria-stabilized zirconia in a tetragonal phase arranged on the bond coat and directly exposed to the hot gas path, a density having a homogeneous distribution such that one standard deviation of the mean or median density is not more than +/- 15% of the mean or median density and having a mean bulk value density of 77-97%, a smooth exposed surface having an Ra of 0.5-2 microns, and a plurality of grooves extending from the single layer surface toward the substrate with dimensions of 50 - 80 z/m width and 40-50% depth of the single layer thermal barrier coating.
  • thermal barrier coating includes: a single layer 6-15% wt.% yttria-stabilized zirconia in a tetragonal phase arranged on or above a substrate onto or below which no additional thermal barrier coating is arranged, a density having a homogeneous distribution such that one standard deviation of the mean or median density is not more than +/- 25% of the mean or median density and having a mean bulk value density of greater than 70%, a smooth exposed surface having an Ra of less than 5 microns, and a plurality of grooves extending from the surface toward the substrate with dimensions of 30 - 300 z/m width and 40- 50% depth of the single layer thermal barrier coating.
  • FIG. 1 illustrates a longitudinal cross-sectional view of a gas turbine engine taken along a plane that contains a longitudinal axis or central axis.
  • FIG. 2 illustrates a combustor section of the gas turbine engine.
  • FIG. 3 illustrates a turbine section of the gas turbine engine.
  • FIG. 4 illustrates a hot gas path extending from the combustor section to the turbine section of the gas turbine engine.
  • FIG. 5 illustrates a layered component designed and constructed for prolonged exposure to the hot gas path.
  • FIG. 6 illustrates an alternative layered component designed and constructed for prolonged exposure to the hot gas path having a single layer thermal barrier coating comprising 6-15% wt.% yttria-stabilized zirconia and having a high homogeneous distribution with a mean bulk value density of greater than 87% and a smooth exposed surface having an Ra of less than 5 microns, and also macro spacing with grooves with dimensions of 30 - 300 »m width and 30- 90% depth of the single layer thermal barrier coating.
  • a single layer thermal barrier coating comprising 6-15% wt.% yttria-stabilized zirconia and having a high homogeneous distribution with a mean bulk value density of greater than 87% and a smooth exposed surface having an Ra of less than 5 microns, and also macro spacing with grooves with dimensions of 30 - 300 »m width and 30- 90% depth of the single layer thermal barrier coating.
  • phrases “associated with” and “associated therewith,” as well as derivatives thereof, may mean to include, be included within, interconnect with, contain, be contained within, connect to or with, couple to or with, be communicable with, cooperate with, interleave, juxtapose, be proximate to, be bound to or with, have, have a property of, or the like.
  • any features, methods, steps, components, etc. described with regard to one embodiment are equally applicable to other embodiments absent a specific statement to the contrary.
  • first”, “second”, “third” and so forth may be used herein to refer to various elements, information, functions, or acts, these elements, information, functions, or acts should not be limited by these terms. Rather these numeral adjectives are used to distinguish different elements, information, functions or acts from each other. For example, a first element, information, function, or act could be termed a second element, information, function, or act, and, similarly, a second element, information, function, or act could be termed a first element, information, function, or act, without departing from the scope of the present disclosure.
  • adjacent to may mean: that an element is relatively near to but not in contact with a further element; or that the element is in contact with the further portion, unless the context clearly indicates otherwise.
  • phrase “based on” is intended to mean “based, at least in part, on” unless explicitly stated otherwise. Terms “about” or “substantially” or like terms are intended to cover variations in a value that are within normal industry manufacturing tolerances for that dimension. If no industry standard is available, a variation of twenty percent would fall within the meaning of these terms unless otherwise stated.
  • axial refers to a direction along a longitudinal axis of a gas turbine engine
  • radial refers to a direction perpendicular to the longitudinal axis of the gas turbine engine
  • downstream refers to a direction along a flow direction
  • upstream refers to a direction against the flow direction.
  • FIG. 1 illustrates an example of a gas turbine engine 100 having a compressor section 102, a combustor section 104 associated therewith, and a turbine section 106 associated therewith.
  • the compressor section 102, combustor section 104 and turbine section 106 being supported by a rotor 112 arranged longitudinally along a central axis 114.
  • the compressor section 102 includes multiple compressor stages each including a set of rotating compressor blades 122 and stationary compressor vanes 124.
  • the compressor section 102 is in fluid communication with an upstream atmospheric air inlet 116 which allows the gas turbine engine 100 to draw atmospheric air into the compressor section 102.
  • the compressor section 102 draws in and compresses the atmospheric air and then diffuses the compressed air for delivery to the combustor section 104.
  • many other arrangements of the compressor section 102 are possible, such as employing one or more stationary or adjustable guide vanes in lieu of or in addition to the stationary compressor vanes 124, depending on the desired design and characteristics of the gas turbine engine 100.
  • the illustrated combustor section 104 includes a plurality of discrete combustors 118 that each operate to mix the compressed air from the compressor section 102 with a flow of fuel to combust that air-fuel mixture to produce a flow of high temperature, high pressure combustion gas 120 that is then directed into the turbine section 106.
  • combustors 118 that each operate to mix the compressed air from the compressor section 102 with a flow of fuel to combust that air-fuel mixture to produce a flow of high temperature, high pressure combustion gas 120 that is then directed into the turbine section 106.
  • many other arrangements of the combustor section 104 are possible, such as employing a single combustor having multiple burners, depending on the desired design and characteristics of the gas turbine engine 100.
  • the turbine section 106 includes multiple turbine stages each including a set of rotating turbine blades 126 and stationary turbine vanes 128.
  • the turbine stages are axially arranged to successively receive the combustion gas 120 from the compressor section 102 and expand that combustion gas 120 to convert thermal and pressure energy into rotating or mechanical work. If the gas turbine engine 100 is used in power generation or prime mover applications, the turbine section 106 is typically connected to a generator, pump, or other device to be driven. As with the compressor section 102 and with the combustor section 104, other turbine section 106 arrangements are possible depending on the desired design and characteristics of the gas turbine engine 100, such as in aero applications.
  • An exhaust section 108 is positioned downstream of the turbine section 106 and arranged to receive the expanded flow of combustion gas 120 from the final turbine stage in the turbine section 106.
  • the exhaust section 108 is arranged to efficiently direct the axially flowing gas away from the turbine section 106 for efficient operation of the gas turbine engine 100.
  • Many variations and design considerations are possible in the exhaust section 108. As such, the illustrated exhaust section 108 is but one example of those variations.
  • a control system 110 optionally may be coupled to the gas turbine engine 100 and operate to monitor various operating parameters and to control various operations of the gas turbine engine 100.
  • the control system 110 has a micro-processor and includes memory devices and data storage devices for collecting, analyzing, and storing data.
  • the control system 110 may provide output data to various devices including monitors, printers, indicators, and the like that allow users to interface with the control system 110 to provide inputs or adjustments.
  • a user may input a power output set point and the control system 110 may adjust the various control inputs to achieve that power output in an efficient manner.
  • control system 110 may also control various operating parameters including, but not limited to variable inlet guide vane positions, fuel flow rates and pressures, engine speed, valve positions, generator load, and generator excitation.
  • the optional control system 110 may also monitor various parameters to assure that the gas turbine engine 100 is operating properly, such as inlet air temperature, compressor outlet temperature and pressure, combustor outlet temperature, fuel flow rate, generator power output, bearing temperature, and the like.
  • FIG. 2 provides further details of an exemplary combustor section 104 of the gas turbine engine 100.
  • the illustrated combustor section 104 includes a compressed air inlet 202, diffusor 204 and transition 206, as well as the combustors 118.
  • compressed air 208 from the rotating compressor blades 122 and the stationary compressor vanes 124 of the compressor section 102 enters the combustor section 104 through the compressed air inlet 202 that may further include an outlet guide vane 210 arranged upstream of the diffusor 204.
  • the diffusor 204 diffuses the compressed air 208 as it flows into a plurality of discrete combustors 118 that each operate to mix the compressed air 208 with a flow of fuel (e.g. natural gas) and the air-fuel mixture is then ignited in the combustors 118 to produce a flow of high temperature, high pressure combustion gas 120.
  • the hot combustion gas 120 is then directed by the transition 206 along a forward portion of the hot gas path to the operatively connected turbine section 106.
  • the rotor 112 may be at least partially or fully enclosed by a stationary shaft cover 212, as exemplary illustrated. If so configured, gaps 214 may exist between components, such as between the stationary shaft cover 212 and the rotor 112. Such illustrated gaps 214 extend in a radial direction and around and between the relatively movable components. Cooling air can leak at gaps 214 between components which negatively affects performance and efficiency of the gas turbine engine 100.
  • a seal or seal assembly may be disposed between one or more gaps 214 to reduce the cooling air leakage.
  • FIG. 3 provides further details of an exemplary turbine section 106 of the gas turbine engine 100.
  • the illustrated turbine section 106 includes a hot gas inlet 302 leading to multiple turbine stages of rotating turbine blades 126 and stationary turbine vanes 128 along the hot gas path 304 upstream of the exhaust section 108.
  • the turbine section 106 is operatively connected to the combustor section 104 typically via the transition 206, although a variety of other configurations could be used such as directly connecting the combustor to a stationary turbine vane 128 depending on the desired design.
  • the turbine section 106 thereby receives the hot combustion gas 120 from the combustor section 104 along a forward portion of the hot gas path 304.
  • the combustion gas 120 then successively impinges upon the multiple turbine stages of rotating turbine blades 126 and stationary turbine vanes 128, whereby the hot combustion gas 120 is expanded to convert the thermal and pressure energy of the combustion gas 120 into rotating or mechanical work, such as rotor 112 rotation or thrust.
  • the turbine section 106 need not have precisely the four illustrated turbine stages but could have more or fewer turbine stages depending on the particular turbine design and corresponding design variables such as power output, turbine length, maximum hot gas path temperature, cooling mechanisms, vane and blade length, needed work extracted, etc.
  • the turbine section 106 may include gaps between components, for which a seal, seal segment, seal assembly, ring, ring segment, seal ring, basket spring, clip and the like may be optionally used to reduce leakage or more efficiently contain and direct the combustion gas 120 along the hot gas path 304, such as between blades and vanes, between blades/vanes and the rotor, between stages, or between tips and segments or between other hot gas path components.
  • a seal, seal segment, seal assembly, ring, ring segment, seal ring, basket spring, clip and the like may be optionally used to reduce leakage or more efficiently contain and direct the combustion gas 120 along the hot gas path 304, such as between blades and vanes, between blades/vanes and the rotor, between stages, or between tips and segments or between other hot gas path components.
  • a seal, seal segment, seal assembly, ring, ring segment, seal ring, basket spring, clip and the like may be optionally used to reduce leakage or more efficiently contain and direct the combustion gas 120 along the hot gas
  • FIG. 4 schematically illustrates the hot gas path 304 taken by the combustion gas 120 in context of an exemplary combustor section 104 having multiple combustors 118 operatively connected to an exemplary turbine section 106 having multiple turbine stages.
  • the hot gas path 304 begins upon ignition of the combustion gas 120 inside the combustor 118 and axially advances as the hot combustion gas 120 flows from the transition 206 through the hot gas inlet 302 thereby forcefully impinging upon first row vane 404 and first row blade 406 to perform work at the first turbine stage 408.
  • the combustion gas 120 then continues its axial advance along the hot gas path 304 thereby forcefully impinging upon second row vane 410 and second row blade 412 to similarly perform work at the second turbine stage 414.
  • the combustion gas 120 axially advances along the hot gas path 304 to the third turbine stage 420, the third row vane 416 and third row blade 418 similarly extract work from the combustion gas 120.
  • the combustion gas 120 axially advances along the hot gas path 304 to fourth turbine stage 426, where the fourth row vane 422 and fourth row blade 424 extract work from the combustion gas 120.
  • the combustion gas 120 has had sufficient velocity, temperature and work drawn from it such that its temperature and velocity is sufficiently reduced and the hot gas path 304 is no longer sufficiently hot and/or of sufficient pressure to be associated with a hot gas path 304 of an industrial gas turbine engine 100.
  • the hot gas path 304 no longer continues as that term is used herein.
  • the hot gas path 304 may extend beyond a fourth turbine stage 426 so long as if additional useful work can be drawn, such as at additional turbine stages, at the exhaust section 108 or at an exit nozzle, as will be understood by those skilled in the art.
  • combustor section 104 and turbine section 106 components are arranged in direct contact with, adjacent to, or associated with the hot gas path 304, including without limitation, combustor 118, transition 206, first row vane 404, first row blade 406, second row vane 410, second row blade 412, third row vane 416, third row blade 418, fourth row vane 422, fourth row blade 424, seals, seal segments and subcomponents thereof including without limitation, blade and vane tips and ring segments.
  • the combustor section 104 and turbine section 106 further includes other components that are in more attenuated indirect (e.g. thermally conductive) contact or otherwise associated with the hot gas path 304 including without limitation, rotor 112, shaft cover 212 and the like.
  • FIG. 5 a cross sectional view of a layered component 500 is shown.
  • the illustrated layered component 500 is designed and constructed for routine prolonged exposure to the hot gas path 304 of a gas turbine engine 100 which reaches temperatures of approximately 1250 degrees Celsius, as well as for direct or indirect exposure to other portions of the hot gas path 304 generally above 1000 degrees Celsius.
  • the illustrated layered component 500 can also be designed and constructed for next generation gas turbine engines 100 having temperatures along portions of the hot gas path 304 that approach 1400 degrees Celsius and potentially even exceed 1700 degrees Celsius.
  • the layered component 500 includes a substrate 502, typically made of a high temperature resistant superalloy.
  • the superalloy-based substrate 502 may often also include internal cooling passages terminating on the component outer surface for passage of coolant fluid to cool the surfaces exposed to the hot gas path.
  • the layered component 500 also includes one or more layers (two exemplarily illustrated) of an even higher temperature resistant ceramic, metallic, or metallic-ceramic thermal barrier coating 506 applied to or above the surface or selected surfaces of the substrate 502 to further assist the layered component 500 to withstand prolonged exposure to the high temperature environment.
  • the layered component 500 may further include one or more intermediary layers (one exemplarily illustrated) of a bond coat 504 applied between the substrate 502 and the thermal barrier coating 506.
  • the structural advantages of the superalloy-based substrate 502 is enhanced by the thermal resistant advantages of the thermal barrier coating 506 and optional intermediate bond coat 504 for prolonged exposure to the hot gas path 304.
  • the layered component 500 nonetheless may be prone to thermal and mechanical failure e.g. cracking due to the extreme heat and prolonged exposure to the hot aggressive combustion gas 120. This cracking, at times, can lead to spallation, i.e., the separation of the insulative thermal barrier coating 506 from the underlying superalloy substrate 502. Such cracking and spallation can increase the temperatures of the exposed underlying substrate 502 significantly, resulting in premature damage and ultimately failure of the layered component 500.
  • the illustrated layered component 500 is advantageously embodied as any one or more of the following hot gas path 304 exposed components: combustor 118, transition 206, first row vane 404, first row blade 406, second row vane 410, second row blade 412, third row vane 416, third row blade 418, fourth row vane 422, fourth row blade 424, seals, seal segments and subcomponents thereof including without limitation, blade and vane tips and ring segments, and/or other components that are in more attenuated indirect (e.g. thermally conductive) contact or otherwise associated with the hot gas path 304 including without limitation, rotor 112, shaft cover 212 and the like.
  • the layered component 500 may be embodied as other components or as a layered component 500 that is exposed to other high temperature environments such as furnaces, combustions sources and exhausts, and the like.
  • the substrate 502 can be made of any of a variety of suitable materials such as superalloys including without limitation: Alloy 247, IN 718, IN 939, Rene80, PWA 1484, PWA 1483, AMH282, Stalll 5 and the like.
  • suitable materials for the substrate 502 include ceramic matrix composites including without limitation, oxide-oxides and nonoxide-nonoxides such as SiC-SiC. Additionally, solely ceramic materials such as those including silica may be appropriately used as well as ceramic-metallic materials such as alumina.
  • the thermal barrier coating 506 can be made of any of a variety of suitable materials such as those that are ceramic based including without limitation 6-15% wt.% yttria-stabilized zirconia, 30-50% wt.% yttria- stabilized zirconia, pyrochlores such as Gd2Zr2O7 and Hf2Zr2O7, as well as bilayer and multilayer YSZ/pyrochlore thermal barrier coating 506.
  • the option bond coat 504 may function as an intermediary layer between the substrate 502 and the thermal barrier coating 506 as part of the overall tough thermal barrier system, as illustrated, in order to improve adhesion of the thermal barrier coating 506 to the substrate 502.
  • the exemplarily illustrated bond coat 504 can be made of any of a variety of suitable materials such as ceramics, metals and ceramic-metals, including without limitation MCrAlY where M denotes a metal such as nickel, cobalt, iron or mixtures thereof, Cr denotes chromium, Al denotes aluminum and Y denotes yttrium, and or other suitable bond coat 504 materials such as alumina.
  • FIG. 6 illustrates another layered component 600 generally similar in design, construction and material composition as the layered component 500, with differences described below.
  • the alternative layered component 600 includes a substrate 502, preferably made of a metallic material and more preferably made of superalloy material; an optional bond coat 504 preferably made of a ceramic-metallic material and more preferably made of a MCrAlY or alumina material; and an alternative thermal barrier coating 602, preferably made of a ceramic material and more preferably made of 6-15% wt.% yttria-stabilized zirconia material within the range for the tetragonal phase version of yttria-stabilized zirconia and most preferably made of a 6-10% wt.% yttria-stabilized zirconia material.
  • the alternative thermal barrier coating 602 advantageously comprises up to four features (that is, not all four features are required to be utilized in order to practice the invention as claimed) that enable the desirous properties of the illustrated alternative thermal barrier coating 602.
  • the first feature involves a single layer 604 of 6-15% wt. % yttria-stabilized zirconia within the range for the tetragonal phase version of yttria-stabilized zirconia, preferably 6-10% wt.%, yttria-stabilized zirconia within the range for the tetragonal phase version of yttria- stabilized zirconia.
  • the single layer 604 does not include or have an adjacent or associated with upper or lower layer of another ceramic thermal barrier coating 506 material such as a pyrochlore or other appreciably different wt. % yttria-stabilized zirconia (e.g. 30-50%).
  • the single layer 604 advantageously provides improved alternative thermal barrier coating 602 surface adhesion to the bond coat 504 (if used, as illustrated) or substrate 502 (if bond coat 504 is not used) and improved overall material strength of the alternative layered component 600.
  • the single layer 604 yttria-stabilized zirconia material can be made by mixing zirconium oxide ZrCh with a predetermined concentration of yttrium oxide Y2O3. Stabilizing the single layer 604 in the tetragonal phase provides superior material strength e.g. fracture resistance and overall toughness relative to a cubic phase.
  • Stabilizing the single layer 604 in the tetragonal phase provides superior thermal resistance relative to a monoclinic phase.
  • the single layer 604 is advantageously stabilized in the tetragonal phase such that it does not transform into a monoclinic or cubic phase during the majority of (i.e. greater than 50%), preferably substantiality of (i.e. greater than 80%), most preferably entirety of (i.e. 100%), the operational life of the alternative layered component 600.
  • the operational life of the alternative layered component 600 begins when the alternative layered component 600 completes its first fire within the gas turbine engine 100 and ends when the alternative layered component 600 begins its final ramp-down within the same gas turbine engine 100, inclusive of the significantly temperature varying intermediate gas turbine engine 100 startup and shutdown activities. Stabilizing the single layer 604 in one phase during the operational life of the alternative layered component 600 inhibits degradation of the alternative thermal barrier coating 602.
  • the second feature involves a density that is substantially uniform. That is, the alternative thermal barrier coating 602 features a density that has a homogeneous distribution 606.
  • the homogeneous distribution 606 has no systemic variation in density through the thickness of the alternative thermal barrier coating 602.
  • the lower portion of the alternative thermal barrier coating 602 is not intentionally or unintentionally denser than the upper portion of the alternative thermal barrier coating 602.
  • the alternative thermal barrier coating 602 is not designed or applied such that the density is variable.
  • spatial variance in density is generally unintended and random.
  • the homogeneous distribution 606 is such that one standard deviation ( I o) of the mean or median density of the alternative thermal barrier coating 602 as measured by total mass divided by total volume is not more than +/- 25% of the mean or median density, and preferably within +/- 15% of the mean or median density, and more preferably within +/- 10% of the mean or median density.
  • the density advantageously has a mean bulk value density 608 (including globular pores, inter splat interfaces and microcracks) of greater than 70% and preferably 77-97%.
  • the mean bulk value density 608 has a porosity including globular pores of under 30% and preferably 3-23%.
  • the third feature involves a smooth exposed surface 610. That is, the alternative thermal barrier coating 602 features an applied as-sprayed condition having an Ra of less than 8 microns, and preferably 2-5 microns.
  • the smooth exposed surface 610 may be advantageously further smoothed in any suitable manner such as abrading or polishing to attain a finished aftersprayed condition having an Ra of less than 5 microns and advantageously 0.5-2 microns. Utilization of a smooth exposed surface 610 advantageously provides the alternative thermal barrier coating 602 with improved aerodynamic efficiency and impact tolerance.
  • the fourth feature involves macro spacing 612 along portions of the alternative thermal barrier coating 602 surface.
  • the illustrated macro spacing 612 is shown as a plurality of spaced and dimensioned valleys or grooves 614 extending from the smooth exposed surface 610 toward the substrate 502 having dimensions of 30 - 300 »m width and 30- 90% of the depth of the alternative thermal barrier coating 602.
  • the macro spacing 612 need not be uniformly spaced or dimensioned, depending on the particular alternative thermal barrier coating 602 or alternative layered component 600.
  • a U or V-shaped macro spacing 612 groove 614 with a width of 50-80 »m and a depth of 40-50% may be used.
  • the macro spacing 612 groove 614 also need not be of any particular or uniform spacing, pattern or geometric shape but rather may be of any suitable spacing, pattern or two or three dimensional geometric or non-geometric shape, such as V-shaped, U-shaped, curved-cut, square, triangular, rectangular, hexagonal, pyramidal, frusto-conical, honeycomb, combinations thereof and the like, depending on the particular alternative thermal barrier coating 602 or alternative layered component 600.
  • the macro spacing 612 advantageously provides the alternative thermal barrier coating 602 with well-placed engineered strain tolerance that assists in thermal resistance particularly when utilized with one or more of the above features.
  • the alternative thermal barrier coating 602 can be applied directly onto the optional bond coat 504 or substrate 502 by any suitable means to advantageously achieve a coating thickness typically of between 0.3 - 4 mm, such as by the known and relatively inexpensive air plasma spraying technique or the less commonly used and relatively expensive but higher density producing physical vapor deposition technique.
  • each of the four above described advantageous features of the alternative thermal barrier coating 602 can be mixed and matched, and modified or adapted within their above identified quantified ranges, in order to accommodate an intended design or configuration of a particular alternative thermal barrier coating 602, which in turn may depend on an intended design or configuration of a particular layered component 500, 600 or design and characteristics of the gas turbine engine 100.
  • the proposed thermal barrier coating 506, alternative thermal barrier coating 602 and layered components 500, 600 provide a simple, cost effective and tough layer for prolonged exposure to a hot gas path 304 of a gas turbine engine 100. This allows for hot gas path 304 components to survive prolonged exposure to the increased hot gas path 304 temperatures of intended future gas turbine engines 100.
  • thermal barrier coating 506 alternative thermal barrier coating 602 and layered components 500, 600 of the present disclosure has been described in detail, those skilled in the art will understand that various changes, substitutions, variations, and improvements disclosed herein may be made without departing from the spirit and scope of the disclosure in its broadest form.

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  • Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Materials Engineering (AREA)
  • Ceramic Engineering (AREA)
  • Combustion & Propulsion (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)

Abstract

Une turbine à gaz comprend une section compresseur, une section chambre de combustion, une section turbine, un trajet de gaz chaud et un élément stratifié (600) comprenant un substrat (502) et un revêtement formant barrière thermique (TBC) (602) présentant : une unique couche de 6 à 15 % en poids de zircone stabilisée par de l'oxyde d'yttrium en phase tétragonale disposée sur ou au-dessus du substrat, l'unique couche ne possédant pas une couche supérieure ou inférieure adjacente d'un second TBC, une masse volumique ayant une distribution homogène de telle sorte qu'un écart type de la masse volumique moyenne ou médiane n'est pas supérieur à +/- 25 % de la masse volumique moyenne ou médiane, et ayant une masse volumique en valeur apparente moyenne supérieure à 70 %, une surface exposée lisse ayant une Ra inférieure à 5 microns, et une pluralité de rainures s'étendant depuis la surface de l'unique couche en direction du substrat avec des dimensions de 30 à 300 µm et une profondeur de 30 à 90 % du TBC.
PCT/US2020/070411 2020-08-14 2020-08-14 Revêtement formant barrière thermique, élément stratifié comprenant un revêtement formant barrière thermique et turbine à gaz comprenant un élément stratifié WO2022035460A1 (fr)

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Citations (3)

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Publication number Priority date Publication date Assignee Title
US20060151856A1 (en) * 2004-12-14 2006-07-13 Taiji Torigoe Thermal barrier coating material, thermal barrier member, and member coated with thermal barrier and method for manufacturing the same
US20070274837A1 (en) * 2006-05-26 2007-11-29 Thomas Alan Taylor Blade tip coatings
JP2012224893A (ja) * 2011-04-18 2012-11-15 Mitsubishi Heavy Ind Ltd 耐熱部材の補修方法、補修耐熱部材

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Publication number Priority date Publication date Assignee Title
US20060151856A1 (en) * 2004-12-14 2006-07-13 Taiji Torigoe Thermal barrier coating material, thermal barrier member, and member coated with thermal barrier and method for manufacturing the same
US20070274837A1 (en) * 2006-05-26 2007-11-29 Thomas Alan Taylor Blade tip coatings
JP2012224893A (ja) * 2011-04-18 2012-11-15 Mitsubishi Heavy Ind Ltd 耐熱部材の補修方法、補修耐熱部材

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