WO2016133580A1 - Revêtement de barrière thermique pour composant de turbine avec éléments de surface verticalement alignés et éléments de rainure à bifurcations multiples - Google Patents

Revêtement de barrière thermique pour composant de turbine avec éléments de surface verticalement alignés et éléments de rainure à bifurcations multiples Download PDF

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Publication number
WO2016133580A1
WO2016133580A1 PCT/US2015/064420 US2015064420W WO2016133580A1 WO 2016133580 A1 WO2016133580 A1 WO 2016133580A1 US 2015064420 W US2015064420 W US 2015064420W WO 2016133580 A1 WO2016133580 A1 WO 2016133580A1
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WO
WIPO (PCT)
Prior art keywords
tbc
layer
groove
component
pattern
Prior art date
Application number
PCT/US2015/064420
Other languages
English (en)
Inventor
Kai Kadau
Jose Antonio PASCUAL-GUTIERREZ
Atin SHARMA
Original Assignee
Siemens Aktiengesellschaft
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
Priority claimed from PCT/US2015/016318 external-priority patent/WO2015130526A2/fr
Application filed by Siemens Aktiengesellschaft, Siemens Energy, Inc. filed Critical Siemens Aktiengesellschaft
Priority to PCT/US2015/064420 priority Critical patent/WO2016133580A1/fr
Priority to CN201580076436.1A priority patent/CN107429573A/zh
Priority to EP15820362.0A priority patent/EP3259453A1/fr
Priority to US15/547,709 priority patent/US20180066527A1/en
Publication of WO2016133580A1 publication Critical patent/WO2016133580A1/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
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D11/00—Preventing or minimising internal leakage of working-fluid, e.g. between stages
    • F01D11/08—Preventing or minimising internal leakage of working-fluid, e.g. between stages for sealing space between rotor blade tips and stator
    • 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
    • F01D11/00—Preventing or minimising internal leakage of working-fluid, e.g. between stages
    • F01D11/08—Preventing or minimising internal leakage of working-fluid, e.g. between stages for sealing space between rotor blade tips and stator
    • F01D11/12—Preventing or minimising internal leakage of working-fluid, e.g. between stages for sealing space between rotor blade tips and stator using a rubstrip, e.g. erodible. deformable or resiliently-biased part
    • 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
    • F01D5/14—Form or construction
    • 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
    • F01D5/14—Form or construction
    • F01D5/18—Hollow blades, i.e. blades with cooling or heating channels or cavities; Heating, heat-insulating or cooling means on blades
    • 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
    • F01D5/14—Form or construction
    • F01D5/18—Hollow blades, i.e. blades with cooling or heating channels or cavities; Heating, heat-insulating or cooling means on blades
    • F01D5/186—Film cooling
    • 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

Definitions

  • the invention relates to combustion or steam turbine engines having thermal barrier coating (“TBC”) layers on its component surfaces, such as blades, vanes, ring segments, or transitions, which are exposed to heated working fluids, such as combustion gasses or high-pressure steam, including individual sub components that incorporate such thermal barrier coatings.
  • TBC thermal barrier coating
  • the invention also relates to methods for reducing crack propagation or spallation damage to such component TBC layers that are often caused by engine thermal cycling or foreign object damage (“FOD”). More particularly, various embodiments described herein relate to the formation of planform patterns of engineered multifurcated groove features ("EGFs”) within the outer surface of the TBC, which are vertically aligned with engineered surface features (“ESFs”) that project upwardly from the component.
  • EGFs include a planform pattern of overlying vertices, which are respectively in vertical alignment with an underlying corresponding ESF. At least three respective groove segments within the EGF pattern converge at each respective overlying vertex in a
  • each converging groove segment has at least two other (i.e., bifurcated) adjoining converging groove segments at each overlying vertex.
  • the vertically aligned ESFs and furcated EGFs localize thermal stress or foreign object damage (FOD) induced crack propagation within the TBC that might otherwise allow excessive TBC spallation and subsequent thermal exposure damage to the turbine component underlying substrate.
  • FOD foreign object damage
  • Known turbine engines including gas/combustion turbine engines and steam turbine engines, incorporate shaft-mounted turbine blades circumferentially circumscribed by a turbine casing or housing.
  • hot combustion gasses flow in a combustion path that initiates within a combustor and are directed through a generally tubular transition into a turbine section.
  • a forward or Row 1 vane directs the combustion gasses past successive alternating rows of turbine blades and vanes. Hot combustion gas striking the turbine blades cause blade rotation, thereby converting thermal energy within the hot gasses to mechanical work, which is available for powering rotating machinery, such as an electrical generator.
  • Engine internal components within the hot combustion gas path are exposed to combustion temperatures approximately well over 1000 degrees Celsius (1832 degrees Fahrenheit).
  • the engine internal components within the combustion path such as for example combustion section transitions, vanes and blades are often constructed of high temperature resistant superalloys. Blades and vanes often include cooling passages terminating in cooling holes on component outer surface, for passage of coolant fluid into the combustion path.
  • Turbine engine internal components often incorporate a thermal barrier coat or coating (“TBC”) of metal-ceramic material that is applied directly to the external surface of the component substrate surface or over an intermediate metallic bond coat (“BC”) that was previously applied to the substrate surface.
  • TBC thermal barrier coat or coating
  • the TBC provides a thermal insulating layer over the component substrate, which reduces the substrate temperature. Combination of TBC application along with cooling passages in the component further lowers the substrate temperature.
  • OTBC outer thermal barrier coating
  • TBC and OTBC are used interchangeably herein when referring to general material properties of the coatings proximate to the coating outer surface that contacts hot working gas in the engine.
  • the outer surface that contacts hot working gas it will be the outer surface of the TBC, in single layer embodiments, or correspondly, the outer surface of the OTBC in multi-layer embodiments.
  • TBC layer loss/delamination negatively affect the TBC layer's structural integrity and potentially lead to its spallation (i.e., separation of the TBC insulative material from the turbine component).
  • spallation i.e., separation of the TBC insulative material from the turbine component.
  • vertical cracks developing within the TBC layer can propagate to the TBC/substrate interface, and then spread horizontally.
  • horizontally oriented cracks can originate within the TBC layer or proximal the TBC/substrate interface.
  • Such cracking loss of TBC structural integrity can lead to further, premature damage to the underlying component substrate.
  • the TBC layer breaks away from underlying substrate the latter loses its protective thermal layer coating.
  • the hot combustion gasses will erode or otherwise damage the exposed component substrate surface, potentially reducing engine operational service life.
  • the TBC layer on engine components is also susceptible to foreign object damage ("FOD") as contaminant particles within the hot combustion gasses strike the relatively brittle TBC material.
  • FOD foreign object damage
  • a foreign object impact can crack the TBC surface, ultimately causing spallation loss of surface integrity that is analogous to a road pothole. Once foreign object impact spalls of a portion off the TBC layer, the remaining TBC material is susceptible to structural crack propagation and/or further spalling of the insulative layer.
  • CMAS calcium, magnesium, aluminum, and silicon
  • intermediate metallic bond coat (BC) layers have been applied directly over the substrate.
  • Structural surface properties and/or profile of the substrate or BC interface to the TBC have also been modified from a flat, bare surface.
  • Some known substrate and/or BC surface modifications e.g., so-called "rough bond coats" or RBCs
  • RBCs have included roughening the surface by ablation or other blasting, thermal spray deposit or the like.
  • the BC or substrate surface has been photoresist or laser etched to include surface features approximately a few microns ( ⁇ ) in height and spacing width across the surface planform.
  • Features have been formed directly on the substrate surface of turbine blade tips to mitigate stress experienced in blade tip coatings.
  • Rough bond coats have been thermally sprayed to leave porous surfaces of a few micron-sized features.
  • TBC layers have been applied by locally varying homogeneity of the applied
  • a weakened zone has been created in the TBC layer corresponding to a known or likely stress concentration zone, so that any cracks developing therein are propagated in a desired direction to minimize overall structural damage to the TBC layer.
  • Various embodiments of turbine component construction and methods for making turbine components help preserve turbine component thermal barrier coating (“TBC”) layer structural integrity during turbine engine operation.
  • TBC turbine component thermal barrier coating
  • ESFs engineered surface features
  • the ESFs function as walls or barriers that contain or isolate cracks in the TBC layer, inhibiting additional crack propagation within that layer or delamination from adjoining coupled layers.
  • the ESFs and vertices of converging EGFs are vertically aligned.
  • engineered groove features are cut and formed in the TBC layer through the outer surface thereof, such as by laser, water jet, or machining, into a previously formed TBC layer.
  • the EGFs functioning as the equivalent of a fire line that prevents a fire from spreading across a void or gap in combustible material— stop further crack propagation in the TBC layer across the groove to other zones in the TBC layer.
  • EGFs in some embodiments are aligned with stress zones that are susceptible to development of cracks during engine operation. In such embodiments, formation of a groove in the stress zone removes material that possibly or likely will form a stress crack during engine operation.
  • EGFs are formed in convenient two dimensional or polygonal planform patterns into the TBC layer.
  • the EGFs localize thermal stress or foreign object damage (FOD) induced crack propagation within the TBC that might otherwise allow excessive TBC spallation and subsequent thermal exposure damage to the turbine component underlying substrate.
  • FOD foreign object damage
  • a given TBC surface area that has developed one or more stress cracks is isolated from non-cracked portions that are outside of the EGFs. Therefore, if the cracked portion isolated by one or more EGFs spalls from the component the remaining TBC surface outside the crack containing grooves will not spall off because of the contained crack(s).
  • spallation of cracked TBC material that is constrained within ESFs and/or EGFs leaves a partial underlying TBC layer that is analogous to a road pothole.
  • the underlying TBC material that forms the floor or base of the "pot hole” provides continuing thermal protection for the turbine engine component underlying substrate.
  • the ESFs have planform patterns of multifurcated groove segments that converge in vertices.
  • the multifurcated, groove geometry is useful for arresting crack propagation in the TBC, whether the crack inducing stress in the TBC is caused by thermo-mechanical stress, induced by heating transients, or foreign object damage (FOD) impact mechanical stress.
  • FOD foreign object damage
  • embodiments of the invention described herein feature combustion turbine engine components, having a heat insulating outer surface for exposure to combustion gas, such as blade, vane, transition, or ring segment abradable components.
  • the component includes a metallic substrate having a substrate surface, and an anchoring layer built upon the substrate surface.
  • a planform pattern of engineered surface features (ESFs) is formed in and projects from the anchoring layer.
  • a thermally sprayed or vapor deposited or solution/suspension plasma sprayed, single- or multi-layer thermal barrier coat (TBC), having a TBC inner surface, is applied over and coupled to the anchoring layer.
  • TBC has a TBC outer surface for exposure to combustion gas.
  • a planform pattern of engineered groove features is cut and formed into the TBC outer surface, penetrating the previously applied TBC layer.
  • the EGFs have groove depth.
  • the EGF pattern defines a planform pattern of overlying vertices, which are respectively in vertical alignment with an underlying corresponding ESF. At least three respective groove segments within the EGF pattern converge at each respective overlying vertex in a multifurcated pattern, so that each converging groove segment has at least two other adjoining converging groove segments at each overlying vertex.
  • FIG. 1 A block diagram illustrating an exemplary combustion turbine engine component
  • FIG. 1 A block diagram illustrating an exemplary combustion turbine engine component
  • FIG. 1 A block diagram illustrating an exemplary combustion turbine engine component
  • FIG. 1 A block diagram illustrating an exemplary combustion turbine engine component
  • FIG. 1 A block diagram illustrating an exemplary combustion turbine engine component
  • FIG. 1 A block diagram illustrating an exemplary combustion turbine engine component
  • FIG. 1 A block diagram illustrating an a combustion turbine engine component
  • FIG. 1 A block diagram illustrating an exemplary combustion turbine engine component
  • TBC thermal barrier coat
  • the TBC has a TBC inner surface that is applied over and coupled to the anchoring layer and a TBC outer surface for exposure to combustion gas.
  • a planform pattern of engineered groove features (EGFs), having groove depths, is cut and formed into the TBC outer surface, penetrating the previously applied TBC layer.
  • the EGF pattern defines a planform pattern of overlying vertices, which are respectively in vertical alignment with an underlying corresponding ESF. At least three respective groove segments within the EGF pattern converge at each respective overlying vertex in a multifurcated pattern, so that each converging groove segment has at least two other adjoining converging groove segments at each overlying vertex.
  • TBC thermal barrier coating
  • the provided component includes a metallic substrate having a substrate surface.
  • An anchoring layer is formed upon the substrate surface.
  • ESFs engineered surface features
  • a thermally sprayed or vapor deposited or solution/suspension plasma sprayed, single- or multi-layer thermal barrier coat (TBC) is applied to the substrate, having a TBC inner surface that is applied over and coupled to the anchoring layer and a TBC outer surface for exposure to combustion gas.
  • a planform pattern of engineered groove features (EGFs), having groove depths, is cut and formed into the TBC outer surface, penetrating the previously applied TBC layer.
  • the EGF pattern defines a planform pattern of overlying vertices, which are respectively in vertical alignment with an underlying corresponding ESF. At least three respective groove segments within the EGF pattern converge at each respective overlying vertex in a multifurcated pattern, so that each converging groove segment has at least two other adjoining converging groove segments at each overlying vertex.
  • the engine including the provided component, is operated, which induces thermal or mechanical stress in the TBC layer during engine thermal cycling or induces mechanical stress in the TBC layer by foreign object impact. If any of the induced stresses generates a crack in the TBC; crack propagation is arrested in the TBC upon intersection with one or more of the EGFs or ESFs.
  • FIG. 1 is a partial axial cross sectional view of a gas or combustion turbine engine incorporating one more exemplary thermal barrier coating ("TBC”) embodiments of the invention
  • FIG. 2 is a detailed cross sectional elevational view of the turbine engine of FIG. 1, showing Row 1 turbine blade and Rows 1 and 2 vanes incorporating one or more exemplary TBC embodiments of the invention;
  • FIG. 3 is a fragmentary view of a turbine component, such as for example a turbine blade, vane or combustion section transition, having an exemplary
  • ESFs engineered surface features
  • BC bond coat
  • FIG. 4 is a fragmentary view of a turbine component, having an exemplary embodiment of ESFs formed directly in the substrate surface with a two layer TBC comprising a lower thermal barrier coat (“LTBC”) applied over the ESFs and an outer thermal barrier coat (“OTBC”) applied over the LTBC;
  • LTBC lower thermal barrier coat
  • OTBC outer thermal barrier coat
  • FIG. 5 is a fragmentary view of an exemplary embodiment of a turbine component having hexagonal planform profile of solid projection ESFs on its substrate surface;
  • FIG. 6 is a cross section of the ESF of FIG. 5;
  • FIG. 7 is a fragmentary view of a turbine component having an exemplary embodiment of a plurality of cylindrical or post-like profile ESFs forming in combination a hexagonal planform pattern on its substrate surface that surround or circumscribes another centrally located post-like ESF;
  • FIG. 8 is a cross section of the ESF of FIG. 7;
  • FIG. 9 is a fragmentary view of a turbine component having an exemplary embodiment of a roughened bond coat ("RBC") layer applied over previously formed ESFs in a lower BC that was previously applied to the component substrate;
  • RBC roughened bond coat
  • FIG. 10 is a fragmentary cross section of a prior art turbine component experiencing vertical and horizontal crack formation in a bi-layer TBC, having a featureless surface BC applied over a similarly featureless surface substrate;
  • FIG. 11 is a fragmentary cross section of a turbine component having an exemplary embodiment of ESFs formed in a LTBC layer, wherein vertical and horizontal crack propagation has been arrested and disrupted by the ESFs;
  • FIG. 12 is a fragmentary perspective view of a turbine component having an exemplary embodiment of engineered groove features ("EGFs") formed in the TBC outer surface;
  • EGFs engineered groove features
  • FIG. 13 is a schematic cross sectional view of the turbine component of FIG. 12 having EGFs formed in the TBC;
  • FIG. 14 is a schematic cross sectional view of the turbine component of FIG. 13 after impact by a foreign object, causing foreign object damage ("FOD") in the TBC, where crack propagation has been arrested along intersections with the EGFs;
  • FOD foreign object damage
  • FIG. 15 is a schematic cross sectional view of the turbine component of FIG. 13 after spallation of an portion of the TBC above the cracks, leaving an intact layer of the TBC below the cracks for continuing thermal insulation of the underlying turbine component substrate;
  • FIG. 16 is a schematic cross sectional view of a turbine component having an exemplary embodiment of a trapezoidal cross section ESF that is anchoring the TBC, with the arrows pointing to stress concentration zones within the TBC;
  • FIG. 17 is a schematic cross sectional view of the turbine component of FIG. 16, in which exemplary embodiments of angled EGFs have been cut into the TBC in alignment with the stress concentration zones in order to mitigate potential stress concentration;
  • FIG. 18 is a schematic cross sectional view of an exemplary embodiment of a turbine component having both ESFs and EGFs;
  • FIG. 19 is a schematic cross sectional view of the turbine component of FIG. 18, in which FOD crack propagation has been constrained by the ESFs and EGFs;
  • FIG. 20 is an exemplary embodiment of EGFs formed in a turbine component TBC outer surface near component cooling holes, in order to arrest propagation of cracks or delamination of the TBC layer in zones surrounding the cooling holes to the surface area on the opposite sides of the grooves;
  • FIG. 21 is a schematic plan view of an exemplary embodiment of a turbine component outer surface OTBC layer, with furcated, EGFs forming hexagon planform patterns therein, with the formed grooves converging at vertices of the hexagons, wherein OTBC layer stress force in the OTBC material along one upstream groove that has induced crack propagation therein is bifurcated at a pair of downstream grooves, thereby arresting further crack propagation in the OTBC material;
  • FIG. 22 is an alternative embodiment of a turbine component outer surface OTBC layer, with furcated, EGFs forming a planform pattern of adjoining hexagons therein, with formed discontinuous grooves converging at a vertices of the hexagon;
  • FIG. 23 is a schematic planform view of an alternative embodiment of a turbine component outer surface OTBC layer, with furcated, EGFs forming varying size and density hexagonal planform patterns across the component surface;
  • FIG. 24 is a schematic planform view of an alternative embodiment of a turbine component outer surface OTBC layer, with furcated, EGFs forming adjoining outer hexagons, which in turn circumscribe furcated EGFs forming nested hexagons and triangular polygons;
  • FIG. 25 is a schematic planform view of an alternative embodiment of a turbine component outer surface OTBC layer, with furcated, EGFs forming an outer hexagon whose converging groove segment vertices are vertically aligned with ESFs projecting from the substrate, the outer hexagon in turn circumscribing furcated EGFs forming triangular polygons that converge at a central vertex over a central ESF;
  • FIG. 26 is a schematic planform view of an alternative embodiment of a turbine component outer surface OTBC layer, with furcated, EGFs forming an outer hexagon whose converging groove segment vertices are vertically aligned with ESFs, the outer hexagon in turn circumscribing furcated EGFs forming adjoining hexagons and trapezoid polygons;
  • FIG. 27 is a schematic planform view of an alternative embodiment of a turbine component outer surface OTBC layer, with furcated, EGFs forming an outer hexagon whose converging groove segment vertices are vertically aligned with ESFs, the outer hexagon in turn circumscribing furcated EGFs forming adjoining hexagons and triangle polygons of different sizes, including a central, nested hexagon vertically aligned with a central ESF; and
  • FIG. 28 is a schematic planform view of an alternative embodiment of a turbine component outer surface OTBC layer, with furcated, EGFs forming an outer hexagon whose converging groove segment vertices are vertically aligned with ESFs, and with other furcated, EGFs forming a grid of smaller hexagons.
  • T thermal barrier coat (TBC) layer thickness
  • Exemplary embodiments of the present invention enhance performance of the thermal barrier coatings (“TBCs”) that are applied to surfaces of turbine engine components, including combustion or gas turbine engines, as well as steam turbine engines.
  • TBCs thermal barrier coatings
  • EGFs engineered groove features
  • the EGFs are formed within the TBC, and more particularly in the outer surface of the TBC.
  • the EGFs are formed in the outer surface of the outer thermal barrier coating (“OTBC”), and selectively are cut to any desired depth, including down to the substrate surface. EGF widths are also selectively varied.
  • the EGFs are formed in furcated planform patterns, meaning multiple grooves converge, or from another alternative relative perspective, diverge in a forked pattern from a common vertex.
  • three grooves converge at a vertex, they are arrayed in a bifurcated pattern, meaning approach of the common vertex from any one of the grooves will diverge into two separate (hence bifurcated) paths away from the common vertex.
  • the furcated EGFs form planform patterns of adjoining hexagons, which share a common groove and two vertices with neighboring adjoining hexagons.
  • the adjoining hexagons are outer hexagons, which respectively circumscribe other planform EGF patterns, such as hexagons, trapezoids, and/or triangles.
  • the furcated EGF planform pattern vertices are vertically aligned with engineered surface features ("ESFs") that project upwardly from the component substrate surface.
  • the multifurcated EGFs isolate and localize thermos-mechanical stress- or foreign object damage ("FOD") -induced crack propagation within the TBC layer, by spreading the stress forces in the OTBC layer adjoining one upstream groove to multiple downstream grooves across their common vertex.
  • FOD foreign object damage
  • the applied upstream thermo-mechanical stress is dissipated or attenuated by the downstream common vertex grooves.
  • the applied upstream thermo-mechanical stress is sufficiently high to fatigue crack the TBC or OTBC material that adjoins the downstream-furcated EGFs, until the stress is transferred to the next set of converging, furcated EGFs in the planform pattern.
  • the transferred stress is in turn furcated in the next furcated EGFs, in cascading fashion.
  • Crack formation is arrested when the furcated stress concentration diminishes sufficiently to be fully attenuated within a downstream zone of the TBC or OTBC material.
  • the furcated EGF pattern with our without vertical alignment of ESFs projecting from the component substrate surface, enables the TBC or OTBC outer surface to self-absorb and dissipate induced thermo-mechanical stress in a minimized surface area.
  • crack propagation and/or resultant spallation is also minimized on the TBC or OTB C outer surface.
  • turbine engines such as the gas or combustion turbine engine 80 include a multi-stage compressor section 82, a combustion section 84, a multi-stage turbine section 86 and an exhaust system 88. Atmospheric pressure intake air is drawn into the compressor section 82 generally in the direction of the flow arrows F along the axial length of the turbine engine 80. The intake air is
  • the engine's rotor and shaft 90 has a plurality of rows of airfoil cross sectional shaped turbine blades 92 terminating in distal blade tips 94 in the compressor 82 and turbine 86 sections.
  • thermal barrier coat
  • each turbine blade 92 has a concave profile high-pressure side 96 and a convex low-pressure side 98. Cooling holes 99 that are formed in the blade 92 facilitate passage of cooling fluid along the blade surface.
  • the high velocity and pressure combustion gas, flowing in the combustion flow direction F imparts rotational motion on the blades 92, spinning the rotor 90. As is well known, some of the mechanical power imparted on the rotor shaft 90 is available for performing useful work.
  • respective upstream vanes 104 and downstream vanes 106 respectively direct upstream combustion gas generally parallel to the incident angle of the leading edge of turbine blade 92 and redirect downstream combustion gas exiting the trailing edge of the blade 92 for a desired entry angle into downstream Row 2 turbine blades (not shown).
  • Cooling holes 105 that are formed in the vanes 104, 106 facilitate passage of cooling fluid along the vane surface. It is noted that the cooling holes 99 and 105 shown in FIG.
  • a typical turbine blade 92 or vane 104, 106 has many more cooling holes distributed about the respective airfoil bodies of much smaller diameter relative to the respective blade or vane total surface area that is exposed to the engine combustion gas.
  • turbine component surfaces that are exposed to combustion gasses are often constructed with a TBC layer for insulation of their underlying substrates.
  • Typical TBC coated surfaces include the turbine blades 92, the vanes 104 and 106, ring segments 120, and related turbine vane carrier surfaces and combustion section transitions 85.
  • the TBC layer for blade 92, vanes 104 and 106, ring segments 120, and transition 85 exposed surfaces are often applied by thermal sprayed or vapor deposition or solution/ suspension plasma spray methods, with a total TBC layer thickness of 300-2000 microns ( ⁇ ).
  • Insulative layers of greater thickness than 1000 microns ( ⁇ ) are often applied to sector shaped turbine blade tip abradable ring segment 110 components (hereafter also referred to generally as an "abradable component”) that line the turbine engine 80 turbine casing 100 in opposed relationship with the blade tips 94.
  • the abradable components 110 have a support surface 112 retained within and coupled to the casing 100 and an insulative abradable substrate 120, which has an outer surface that is in opposed, spaced relationship with the blade tip 94 by a blade tip gap G.
  • the abradable substrate 120 is often constructed of a metallic/ceramic material, similar to the TBC coating materials that are applied to blade 92, vanes 104, 106 and transition 85 combustion gas exposed surfaces.
  • abradable substrate materials have high thermal and thermal erosion resistance and maintain structural integrity at high combustion temperatures.
  • some form of TBC layer is formed over the blade tip abradable component 110 bare underlying metallic support surface substrate 112 for insulative protection, plus the insulative substrate thickness that projects at additional height over the TBC.
  • the ring segment abradable components 110 have a functionally equivalent TBC layer to the TBC layer applied over the turbine transition 85, blade 92 and vanes 104/106.
  • the abradable surface 120 function is analogous to a shoe sole or heel that protects the abradable component support surface substrate 112 from wear and provides an additional layer of thermal protection.
  • Exemplary materials used for blade tip abradable surface ridges/grooves include pyrochlore, cubic or partially stabilized yttria stabilized zirconia. As the abradable surface metallic ceramic materials is often more abrasive than the turbine blade tip 94 material a blade tip gap G is maintained to avoid contact between the two opposed components that might at best cause premature blade tip wear and in worse case circumstances might cause engine damage.
  • the ring segment abradable components 110 are often constructed with a metallic base layer support surface 112, to which is applied a thermally sprayed ceramic/metallic abradable substrate layer of many thousands of microns thickness (i.e., multiples of the typical transition 85 blade 92 or vanes 104/106 TBC layer thickness).
  • a thermally sprayed ceramic/metallic abradable substrate layer of many thousands of microns thickness (i.e., multiples of the typical transition 85 blade 92 or vanes 104/106 TBC layer thickness).
  • the ring segment 120 abradable surface 120 planform and projection profile embodiments described in the related patent applications for which priority is claimed herein include grooves, depressions or ridges in the abradable substrate layer 120 to reduce abradable surface material cross section for potential blade tip 94 wear reduction and for directing combustion airflow in the gap region G.
  • Commercial desire to enhance engine efficiency for fuel conservation has driven smaller blade tip gap G specifications: preferably no more than 2 millimeters and desirably approaching 1 millimeter ( ⁇
  • Some exemplary turbine component embodiments incorporate an anchoring layer of ESFs that aid mechanical interlocking of the TBC layer and aid in isolation of cracks in the TBC layer, so that they do not spread beyond the ESF.
  • ESFs solid ridge and groove projecting surface features as well as micro surface features
  • the ESFs are formed in an anchoring layer that is coupled to an inner surface layer of the TBC layer and they are sized to anchor the TBC layer coating thickness range of 300-2000 microns ( ⁇ ) applied to those components without changing an otherwise generally flat outer surface of the TBC layer that is exposed to combustion gas.
  • the ESFs have heights and three-dimensional planform spacing on the turbine component surface sufficient to provide mechanical anchoring and crack isolation within the total thickness of the TBC layer.
  • the ESFs will be shorter than the total TBC layer thickness but taller than etched or engraved surface features that are allegedly provided to enhance adhesion bonding between the TBC and the adjoining lower layer (e.g., an underlying naked substrate or intermediate BC layer interposed between the naked substrate and the TBC layer).
  • the ESFs have a projection height between approximately two to seventy- five percent (2-75 %) of the TBC layer's total thickness.
  • the ESFs have a projection height of at least approximately thrity-three percent (33%) of the TBC layer's total thickness. In some exemplary embodiments, the ESFs define an aggregate surface area at least twenty percent (20%) greater than an equivalent flat surface area.
  • FIGs. 3 and 4 show exemplary embodiments of ESFs formed in an anchoring layer that is coupled to an inner surface of the TBC layer.
  • the TBC layer 306/326 may comprise multiple layers of TBC material, but will ultimately have at least a TBC with an outer surface for exposure to combustion gas. In FIG.
  • the turbine component 300/320 for example a combustor section transition, a turbine blade or a turbine vane, has a metallic substrate 301 that is protected by an overlying TBC.
  • a BC layer 302 is built upon and applied over the otherwise featureless substrate 301, which incorporates a planform pattern of ESFs 304.
  • ESFs 304 are formed directly in the BC by: (i) known thermal spray of molten particles to build up the surface feature or (ii) known additive layer manufacturing build-up application of the surface feature, such as by 3-D printing, sintering, electron or laser beam deposition or (iii) known ablative removal of substrate material manufacturing processes, defining the feature by portions that were not removed.
  • the ESFs 304 and the rest of the exposed surface of the BC layer 302 may receive further surface treatment, for example surface roughening, micro engraving or photo etching processes to enhance adhesion of the subsequent thermally sprayed TBC layer 306.
  • the ESFs 304 and the remaining exposed surface of the BC layer 302 comprise an anchoring layer for the TBC layer 306.
  • the outer surface of the TBC layer 306 is exposed to combustion gas.
  • turbine component 320 has an anchoring layer construction, where the planform array of ESFs 324 are formed directly in the otherwise featureless substrate 321, by known direct casting or build-up on the substrate surface by thermal spraying, additive layer build up or, alternatively, by known ablative or other mechanical removal of substrate material, manufacturing processes that defines the feature by remaining portions of the substrate that were not removed.
  • the ESFs 324 and the exposed surface of the naked substrate 321 may receive further surface treatment, for example surface roughening, micro engraving or photo etching processes to enhance adhesion of the subsequent thermally sprayed TBC layer 326.
  • the ESFs 324 and the naked substrate surface comprise an anchoring layer for the TBC layer 326 without any intermediate BC layer.
  • a multi-layer TBC 326 is applied over the anchoring layer.
  • the multi-layer TBC layer 326 comprises a lower thermal barrier coat ("LTBC") 327 layer that is coupled to anchoring layer (in some embodiments the LTBC functions as a portion of the anchoring layer) and an outer thermal barrier coat (“OTBC") layer 328 that has an outer surface for exposure to combustion gas. Additional TBC intermediate layers 326 may be applied between the LTBC layer 327 and the OTBC layer 328.
  • a multi-layer TBC layer is applied over any other type of ESFs that have been previously described. For example, while not shown in the figures, a variation of the construction of the turbine component 300 of FIG. 3, with the ESFs 304 formed in the BC layer 302, has a multilayer TBC 306 similar to the TBC layer 326 applied over the ESFs 304.
  • ESF cross sectional profiles, their planform array patterns, and their respective dimensions may be varied during design and manufacture of the turbine component to optimize thermal protection by inhibiting crack formation, crack propagation, and TBC layer spallation.
  • Different exemplary permutations of ESF cross sectional profiles their three-dimensional planform array patterns and their respective dimensions are shown in FIGs.5 -9.
  • ESF height, ESF ridge width, ridge spacing, and groove width between ridges are illustrated.
  • the ESFs are selectively arrayed in three-dimensional planform linear or polygonal patterns. For example, the ESF planform pattern of parallel vertical projections shown in FIGs.
  • the turbine component 340 has, a metallic substrate 341 with ESFs 344 formed therein, comprising a hexagonal planform of dual grooves circumscribing an upper groove.
  • the turbine component 350 has, a metallic substrate 351 with ESFs 354 formed therein, comprising cylindrical pins.
  • the turbine components 340 and 350 are shown without a TBC layer covering the ESFs 344 or 354.
  • the ESFs 344 or 354 are generally repeated over at least a portion of the surface of their respective substrates.
  • the spacing pitch and footprint size of the three-dimensional planform patterns can also be varied locally on the surface topology of the turbine component.
  • the ESFs shown in FIGs. 5-8 are formed directly in their respective substrates, as previously discussed they may be formed in a BC that is applied over a featureless substrate. It is also noted that additional anchoring capability can be achieved by applying a rough bond coat ("RBC") layer over the anchoring layer surface, such as the RBC layer 365 of the turbine component 360 shown in FIG. 9. While the RBC 365 is shown applied over the BC 362 and its ESFs 364, it or other types of BCs 362 can also be applied directly over the component metallic substrate 361.
  • RBC rough bond coat
  • TBC layer-anchoring advantages provided by the ESFs described herein, they also localize TBC layer crack propagation.
  • thermally and/or foreign object induced cracks 389V and 389 H have formed in an outer TBC layer 388 of bi-layer TBC 386.
  • the inner TBC layer 387 usually having different material properties than the outer TBC layer 388, is coupled to a BC layer 382, with the BC layer 382 in turn coupled to the component metallic substrate 381.
  • the right-most vertical crack 389V has penetrated to the interface of the outer TBC 388 and inner 387 TBC layers and is now propagating horizontally as crack 389H.
  • the metallic substrate 391 also has a BC over layer 382 to which is affixed a TBC layer 396.
  • the TBC layer 396 further comprises a lower thermal barrier coating ("LTBC") layer 397 that has ESFs 394 formed therein for interlocking with the outer thermal barrier coat (“OTBC”) layer 398.
  • LTBC thermal barrier coating
  • the LTBC layer 397 with its ESFs 394 effectively functions as the anchoring layer for the OTBC layer 398.
  • the LTBC layer 397 has greater strength and ductility material properties than the OTBC layer 398, while the latter has greater thermal resistivity and brittleness material properties.
  • Some exemplary turbine component embodiments incorporate planform arrays of engineered groove features ("EGFs”), which are formed in the outer surface of the TBC after the TBC layer application. Groove depth and width are selectively varied. In some embodiments grooves cut into some or all thermal barrier coating layers, engineered surface features (ESFs), bond coat (BC) layers, or even into the underlying substrate surface.
  • EGFs groove axes are selectively oriented, at any skew angle relative to the TBC outer surface and extend into the TBC layer.
  • the EGFs Analogous to a firefighter fire line, the EGFs isolate cracks in the TBC layer, so that they do propagate across the boundary of a groove void into other portions of adjoining TBC material. Generally, if a crack in the TBC ultimately results in spallation of material above the crack the EGF array surrounding the crack forms a localized boundary perimeter of the spall site, leaving TBC material outside the boundary intact. Within the spallation zone bounded by the EGFs, damage will be generally limited to loss of material above the EGF groove depth. Thus in many exemplary embodiments EGF depth is limited to less than the total thickness of all TBC layers, so that a volume and depth of intact TBC material remains to provide thermal protection for the local underlying component metallic substrate. In some embodiments, the EGF arrays are combined with ESF arrays to provide additional TBC integrity than either might provide alone.
  • FIGs. 28 and 13 show a turbine component 400 having an underlying metallic substrate 401 onto which is affixed a TBC substrate 402 with an exemplary three-dimensional planform array of orthogonally intersecting engineered groove features EGFs 403, 404 that were formed after TBC layer application .
  • the grooves 403 and 404 are constructed with one or more groove depths Do, groove widths WG, groove spacing SG, and/or polygonal planform array pattern. Pluralities of any of different groove depth, spacing, width, and polygonal planform pattern can be varied locally about the turbine component 400 surface. For example,
  • TBC layer 402 is shown as directly coupled to the substrate 401 intermediate anchoring layer constructions previously described can be substituted in other exemplary embodiments, including one or more of bond coat (“BC”) or lower thermal barrier coat layers (“LTBC").
  • BC bond coat
  • LTBC lower thermal barrier coat layers
  • Exemplary engineered groove feature (“EGF”) crack isolation capabilities are shown in FIGs. 14 and 15, wherein a turbine component 400, such as a combustion section transition 85, a turbine blade 92, or a turbine vane 104/106 sustains foreign object (“FOD”) impact damage, resulting in vertical and horizontal cracks 408H and 408V within its TBC 402 outer surface 405.
  • the EGFs 404 flanking the impact damage stop further crack propagation across the groove void, sparing TBC material outside the groove boundaries from further cascading crack propagation.
  • TBC layer 402 material bounded by the cracks and the cratered floor 406 protects the underlying metallic substrate 401 from further damage.
  • the engineered groove feature (“EGF") embodiments herein form cut or ablated grooves or other voids through the previously formed TBC layer outer surface to a desired depth.
  • the turbine component 410 has an anchoring layer 412 that includes trapezoidal cross sectional profile engineered surface features("ESFs") 414.
  • the arrows in FIG. 17 identify likely sites in the TBC layer 416 for actual or potential thermal or mechanical stress concentration zones ⁇ at the intersecting edges or vertices of the ESF 414 during turbine engine operation. Accordingly, EGFs 418 are cut at an angle along the stress line ⁇ at a skewed groove axis angle into the TBC outer surface. The EGFs 418 are also cut at sufficient depth to intersect the ESF 414 vertices. Stresses induced in the TBC layer 416 on either side of the EGFs 418 do not propagate from one side to the other. The TBC layer 416 on either side of an EGF 418 is free to expand or contract along the groove void, further reducing likelihood of crack generation parallel to the groove.
  • FIGs. 17-19 show additional TBC crack inhibition and isolation advantages afforded by combination of engineered groove features ("EGFs") and engineered surface features ("ESFs").
  • EGFs engineered groove features
  • ESFs engineered surface features
  • FIG. 16 the advantages of relieving actual or potential stress lines ⁇ were achieved by forming the EGF 418 all the way through the TBC 416 depth until it intersected the anchoring layer's ESF 414.
  • the turbine component 420 e.g., turbine blade or vane or transition
  • metallic substrate 421 has a bond coat (“BC”) 422 anchoring layer, which defines engineered surface features (“ESFs”) 424 that are oriented in a three-dimensional planform pattern.
  • BC bond coat
  • ESFs engineered surface features
  • the TBC layer 426 is applied over the anchoring layer and after which another planform three-dimensional pattern of EGFs 428 are cut through the TBC layer outer surface 427 that is exposed to combustion gasses.
  • the EGF 428 planform patterns may differ from the ESF 424 planform patterns. If the same planform pattern is used for both the ESFs and the EGFs, their respective patterns do not necessarily have to be vertically aligned within the TBC layer(s). In other words, the EGFs and ESFs may define separate three-dimensional, independently aligned planform patterns across the component. In some embodiments the ESFs and EGFs, respectively have repeating
  • Patterns may vary locally about the component surface.
  • the EGF 428, planform pattern does not have any specific alignment that repetitively corresponds to the ESF 424 pattern.
  • EGFs 428 is cut into the ESF 424 ridge plateaus and others only cut into the TBC 426 layer.
  • a foreign object (“FO") has impacted the TBC outer surface 427, creating cracks that are arrested by the ESFs 424A, 424B, and the EGFs 428A and 428B that bound or otherwise circumscribe the FO impact zone.
  • the remaining, non-damaged TBC material 426A that remains affixed to the BC anchoring layer 422 at the base of the "pot hole” provides thermal protection to its underlying metallic substrate 421.
  • engineered groove features can be formed in the TBC layer around part of or the entire periphery of turbine component cooling holes or other surface discontinuities, in order to limit delamination of the TBC over layer along the cooling hole or other discontinuity margins in the component substrate.
  • the TBC layer at the extreme margin of the cooling hole can initiate separation from the metallic substrate that can spread laterally/horizontally within the TBC layer away from the hole. Creation of an EGF at a laterally spaced distance from the cooling hole margin—such as at a depth that contacts the anchoring layer or the metallic substrate—limits further delamination beyond the groove.
  • the turbine component 490 for example a turbine blade or a turbine vane, has a plurality of respective cooling holes 99/105 that are fully circumscribed by the linear EGF segments 494 and 496 of turbine component 490 fully or partially circumscribe cooling holes 99/105 from each other. TBC delamination along one or more of the cooling hole 99/105 peripheral margins is arrested at the intersection of the circumscribing EGF segments 494 and 496.
  • hole periphery EGFs is limited to the groove shape and orientation. Underlying substrate, anchoring layer, ESF and any other EGFs are constructed in accordance with prior descriptions previously as described.
  • the engineered groove feature ("EGF") planform pattern embodiments of FIGs. 21-28 incorporate converging groove segments, at least three of which, in repetitive patterns, share a common vertex.
  • each groove terminus at its common vertex furcates, or branches out to at least two other diverging grooves, which is analogous to an upstream water stream splitting into two downstream tributary streams.
  • the flow volume is divided between the two downstream tributaries.
  • the downstream flow volume in either tributary is less than the upstream flow volume.
  • Localized downstream material in the TBC or OTBC absorbs the induced, now bifurcated, or reduced applied stress that crossed the common vertex boundary. If the downstream-localized material has sufficient strength to avoid cracking, any upstream cracking is thereby arrested. If the downstream-localized material cracks, the applied stress (and possibly the crack) propagates in cascading fashion to the next one or more common vertices. Cascading propagation continues until stress is reduced sufficiently to arrest further crack formation.
  • FIG. 21 is illustrative of an exemplary embodiment of furcated, engineered groove features ("EGFs") in the TBC outer surface of a turbine blade, vane, or transition component 500.
  • the EGFs form a hexagonal- or honeycomb-shaped planform pattern of adjoining hexagons 502, respectively having six grooves 504, which terminate in six vertices 505.
  • Each pair of adjoining hexagons 502 shares a common groove segment 504A and a pair of two vertices 505 A, 505B.
  • Each shared common vertex 505 has three converging groove segments 504.
  • the trio of grooves 504 at each shared vertex 505 is oriented at 120 degrees.
  • the three converging grooves (see, e.g., grooves 509, 511 and 512) respectively bifurcate into the other two adjoining grooves (see, e.g., groove 509 bifurcating into grooves 511 and 512).
  • grooves 509, 511 and 512 respectively bifurcate into the other two adjoining grooves.
  • the bifurcated, or in some embodiments multifurcated, groove geometry concept of FIG. 21 is useful for arresting crack propagation in the OTBC or TBC outer surface, whether the crack inducing stress in the TBC is caused by thermo- mechanical stress, induced by heating transients, or foreign object damage ("FOD") impact mechanical stress.
  • crack-inducing stress OA initiated within the boundaries of the hexagons 506 and 507 will either be dissipated by the TBC material volume within those hexagons (i.e., arrested therein), or the stress- induced crack in the TBC material will eventually intersect one or more of the groove segments 511, 512 in the circumscribing hexagonal boundary of hexagon 508.
  • the stress GA propagates within any groove, such as groove 509, it will be either (i) arrested in its entirety before reaching a boundary vertex 510 or (ii) continue propagation GB and GC into the two adjoining downstream groove segments 511 and 512 that share the common vertex 510.
  • the stress GA propagates to two adjoining downstream groove segments 511, 512, the stress is bifurcated by some ratio, so that the resultant absolute stress level GB and GC in each adjoining hexagon (here hexagon 508) bounded by the respective downstream groove segments 511, 512 is lower than the absolute stress level GA in the upstream, transferring hexagons 506 and 507.
  • the EGF groove segments 524 forming the hexagonal planform pattern are discontinuous, and do not converge into a commonly-communicating groove at each vertex 525, unlike the continuously communicating grooves 504 of the hexagonal planform pattern 502 FIG. 21.
  • discontinuous EGF groove segment construction shown in FIG. 22 may be incorporated into any of the EGF embodiments shown and described in connection with any of the other figures herein, including the
  • FIGs. 21 and 23-28 are embodiments of FIGs. 21 and 23-28.
  • the adjoining hexagonal honeycomb patterns have different size and pitch density in different surface regions of the blade, vane, ring segment or transition component 540 TBC or OTBC coating outer surface.
  • the optimal length scale for the suggested structures will depend on the TBC system (i.e., base material, bond coat, and TBC layers), the local temperature differences during the engine operating cycle, and the local topography of the component.
  • the localized pitch and density pattern is optimized for its intended operating conditions. For example, distance between the hexagonal vertices and their converging groove segments might be larger in the blade root or blade platform portions of a turbine blade, as compared to distance on the blade's leading edge.
  • EGF pitch and density are locally tailored to topographic differences, localized thermal stresses, and risk of foreign object damage ("FOD"). Focusing on blade leading edge operating conditions, its relatively large curvature, high exposure to combustion gasses and foreign objects entrained in the combustion gas, and combustion contaminant degradation of the TBC favors higher density, smaller honeycomb patterns, such as those of the rightmost planform pattern 542 in FIG. 23, whereas the blade pressure side surface might favor the intermediate size honeycomb pattern 544 in the central portion of that figure. The relatively larger honeycomb pattern 546 on the leftmost side of FIG. 23 might be suitable for the blade suction side surface and blade platform.
  • EGF groove cross sectional depth and width can be selectively varied locally in different surface regions of the blade, vane, or transition component 550 TBC or OTBC coating outer surface, in order to control stress and crack propagation, as shown in FIG. 24.
  • Polygonal planform patterns are included within the
  • the outer hexagon 560 in the continuous planform pattern circumscribes two nested hexagons: intermediate hexagon 570 and inner hexagon 580.
  • the regions between the respectively nested hexagons 560, 570, 580 are filled with triangular sub regions in the shape of the triangles 590, 600, 610, with each triangle vertex having at least three converging groove segments.
  • Triangle 590 comprises groove segments 592 and common vertices 594.
  • the groove segments 592 that adjoin the outer hexagon 560 are co-extensive with portions of the groove segments 562, while in some locations the common vertices 564 and 594 are co-extensive.
  • the groove segments 592 that adjoin the intermediate hexagon 570 are co-extensive with portions of the groove segments 572, while in some locations the common vertices 574 and 594 are co-extensive.
  • the triangle 600 Moving inwardly within the nested hexagonal patterns, the triangle 600 has three groove segments 602 and common vertices 604. In some locations, the groove segments 602 are co-extensive with adjoining groove segments 572 or 582, which form the respective intermediate hexagons 570 and inner hexagons 580, and the common vertices 604 are in some locations co-extensive with the common vertices 574 or 584.
  • a stress concentration leading to crack formation distributes the stress constrained by the exemplary triangle 610 region to one or more of the vertices 614 or 584.
  • Those vertices have respective downstream-furcated groove segments that form other adjoining triangles 610 or the inner hexagon 580.
  • the crack-inducing stress dissipates as it cascades through the OTBC material downstream of each cascading, successive groove segment 612 or 582.
  • any one or more of the triangle 610 or hexagon 580 polygons is sufficient to cause a localized surface spalling, the spallation surface damage is minimized and constrained by the remaining, undamaged adjoining polygons, such as the triangles 600.
  • individual grooves forming the cascading EGFs have any desired groove dimensions or planform patterns, as previously described herein.
  • the outer hexagons 560 have wider and/or deeper grooves 562 than the inner circumscribed polygons 570, 580, 590 600, or 610.
  • the intermediate hexagon grooves 572 are narrower and/or shallower than the grooves 562, while the grooves 582 are in turn narrower and/or shallower than the grooves 572.
  • any of the grooves 592, 602 and/or 612 in adjoining triangles, which are intermediate and skewed relative to the nested hexagon grooves 562, 572 and/or 582 are shallower and or narrower than those of the aforementioned hexagon grooves.
  • Any of the aforementioned furcated grooves are formed by any manufacturing method previously described herein. The more groove segments that converge at each vertex furcates the upstream stress forces proportionately to the number of those segments. In this way, the stress force transferred to any of the downstream, multifurcated groove segment-bounded OTBC material is lower than the transferred stress force in the upstream, transferring groove segment-bounded OTBC material.
  • thermal barrier coated (“TBC”) blades, vanes, ring segment abradable surfaces, or combustion gas transition components 630 have composite, vertically aligned engineered surface features ("ESFs”) 632, 634 and engineered groove features (“EGFs”) 642, 652, which combine the coating anchoring enhancement properties of the ESFs with the
  • ESFs 632, 634 are constructed in any desired density, cross section footprint, or height, as previously described.
  • a plurality of cylindrical shaped ESFs 632 aligns with vertices 644 of overlying outer hexagon planform pattern 640 EGF groove segments 642.
  • the ESFs 632 have similar construction to the ESFs 354 of FIGs. 7 and 8.
  • the ESFs are formed in a hexagonal pattern as the ESFs 344 of FIGs. 5 and 6.
  • the respective turbine vane, blade, ring segment abradable surface, or combustion gas transition component has a planform pattern of adjoining, respective outer hexagons 670, or 690, or 710, whose respective vertices 674, or 694, or 714 are oriented in vertical alignment within the planform of the respective cylindrical EGF 676, or 696, or 716 footprints.
  • 26-28 patterns of smaller polygonal hexagons 680, 700, or 702, or 720; half-hexagon-shaped trapezoids 682, or one-third-hexagon-shaped trapezoids 705; or triangles 704 are circumscribed by the respective outer hexagons 670, or 690, or 710. Spallation of any of the smaller polygons leaves the remaining smaller polygons covering and protecting the component.
  • the smaller polygons are combinations of hexagons 700, 702, triangles 704, and trapezoids 705 that are circumscribed by the larger outer hexagon 690.
  • the larger circumscribing outer hexagons 640, 670, 690 and/or 710 of FIGs. 25-28 adjoin other similarly sized hexagons, or they abut against smaller hexagons, as in the planform local pattern of FIG. 23.
  • planform patterns of FIGs. 25-28 are discontinuous clusters of the outer hexagons that are arrayed in uniform or varying pitch and size patterns, or individual stand-alone outer hexagons 640, 670, 690, and/or 710.
  • the furcated groove EGF patterns of FIGs. 25-28 further define within each outer hexagon a planform pattern of adjoining inner polygons. Adjoining inner polygons respectively share at least one common inner polygonal vertex, and each is respectively fully circumscribed within a corresponding respective outer hexagon 670, 690, or 710.
  • At least three respective furcated groove segments within the EGF pattern converge at each respective outer hexagonal or inner polygonal vertex in a bifurcated pattern, so that each converging groove segment has at least two other adjoining converging groove segments.
  • a larger number of converging grooves in the planform pattern increases furcation of the transferred stress forces.
  • the larger hexagon EGFs with or without underlying, vertically aligned ESFs circumscribe thermal or mechanical stress concentration zones within the outer thermal barrier coating ("OTBC"), such as around cooling holes, analogous to the cooling hole groove embodiment of FIG. 20.
  • OTBC outer thermal barrier coating
  • the EGFs have a skewed groove axis, analogous to the grooves 418 of FIG. 17.
  • the stress is attenuated and dissipated at each successive adjoining polygon as the stress force is furcated successively at each groove juncture vertex. Further crack propagation is arrested within one or more of successive inner polygons through which the crack propagates at its intersection with one or more of the groove segments defining the respective polygon, or upon its intersection with one or more of the groove segments defining a circumscribing hexagon. The crack propagates to other adjoining, circumscribing hexagons, if the crack is not arrested in the initially damaged hexagon.
  • the aggregate thermally sprayed TBC layer of any turbine component embodiment described herein may have different local material properties laterally across the component surface or within the TBC layer thickness dimension.
  • one or more separately applied TBC layers closest to the anchoring layer may have greater strength, ductility, toughness and elastic modulus material properties than layers closer to the component outer surface but the higher level layers may have greater thermal resistivity and brittleness material properties.
  • a multi-layer TBC embodiment 326 is shown in FIG. 4.
  • a graded TBC layer construction can be formed by selectively varying constituent materials used to form the TBC layer during a continuous thermal spraying process.
  • a calcium-magnesium-alumino-silicate (“CMAS”), or other contaminant deposit-resistant layer, is applied over TBC outer surface, for inhibiting adhesion of contaminant deposits to the TBC outer surface.
  • CMAS calcium-magnesium-alumino-silicate
  • Undesirable contaminant deposits can alter material properties of the TBC layer and decrease aerodynamic boundary conditions along the component surface.
  • CMAS-resistant layer is applied over and infiltrates EGF grooves that are formed in the TBC outer surface layer it enhances aerodynamic boundary conditions by forming a relatively smoother TBC outer surface and inhibits debris accumulation within the grooves.
  • Exemplary material compositions for thermal barrier coat (“TBC”) layers include yttria-stabilized zirconia, rare-earth stabilized zirconia with a pyrochlore structure, rare-earth stabilized fully stabilized cubic structure, or complex oxide crystal structures such as magnetoplumbite or perovskite or defective crystal structures.
  • Other exemplary TBC material compositions include multi-element-doped oxides with high defect concentrations.
  • Examples of CMAS retardant compositions include alumina, yttrium aluminum oxide garnet, slurry deposited/infiltrated highly porous TBC materials (the same materials that are utilized for OTBC or LTBC compositions), and porous aluminum oxidized to form porous alumina.

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  • General Engineering & Computer Science (AREA)
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  • Materials Engineering (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)

Abstract

Composants pour moteur à turbine (80), tels que des pales (92), des aubes (104, 106), des surfaces abradables (120) de segments de bague (110), ou des transitions (85), comportant des éléments de surface mis au point (ESF) alignés verticalement (632, 634) et des éléments de rainure mis au point (EGF) bifurqués (642, 652). Un motif de forme plane d'EGF (642, 652) est découpé dans la surface extérieure du revêtement de barrière thermique (TBC) du composant. Le motif EGF comprend un motif de forme plane de sommets superposés (644) respectivement en alignement vertical sur un ESF (632, 634) correspondant sous-jacent. Au moins trois segments de rainure respectifs (642, 652, 642) dans le motif EGF (640) convergent au niveau de chaque sommet respectif (644) selon un motif à multiples bifurcations, de sorte que les contraintes induites par des fissures sont atténuées en cascade, quand la contrainte (σA) est bifurquée (σB, σC) à chaque jointure de sommet.
PCT/US2015/064420 2014-02-25 2015-12-08 Revêtement de barrière thermique pour composant de turbine avec éléments de surface verticalement alignés et éléments de rainure à bifurcations multiples WO2016133580A1 (fr)

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PCT/US2015/064420 WO2016133580A1 (fr) 2015-02-18 2015-12-08 Revêtement de barrière thermique pour composant de turbine avec éléments de surface verticalement alignés et éléments de rainure à bifurcations multiples
CN201580076436.1A CN107429573A (zh) 2014-02-25 2015-12-08 具有竖直对准的设计表面和多分叉槽特征的涡轮机部件热障涂层
EP15820362.0A EP3259453A1 (fr) 2015-02-18 2015-12-08 Revêtement de barrière thermique pour composant de turbine avec éléments de surface verticalement alignés et éléments de rainure à bifurcations multiples
US15/547,709 US20180066527A1 (en) 2015-02-18 2015-12-08 Turbine component thermal barrier coating with vertically aligned, engineered surface and multifurcated groove features

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PCT/US2015/016318 WO2015130526A2 (fr) 2014-02-25 2015-02-18 Revêtement formant une barrière thermique pour pièce de turbine présentant des éléments de rainure usinés pour isoler les fissures
USPCT/US2015/016331 2015-02-18
USPCT/US2015/016318 2015-02-18
PCT/US2015/016331 WO2015130528A1 (fr) 2014-02-25 2015-02-18 Revêtement de barrière thermique de composant de turbine avec éléments de surface usinés d'isolation contre les fissures
PCT/US2015/064420 WO2016133580A1 (fr) 2015-02-18 2015-12-08 Revêtement de barrière thermique pour composant de turbine avec éléments de surface verticalement alignés et éléments de rainure à bifurcations multiples

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WO2011085376A1 (fr) * 2010-01-11 2011-07-14 Rolls-Royce Corporation Éléments d'atténuation d'une contrainte thermique ou mécanique sur un revêtement anticorrosion protégeant de l'environnement

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EP3176378A1 (fr) * 2015-12-04 2017-06-07 United Technologies Corporation Revêtement barrière thermique à pouvoir d'adhésion amélioré
US10731482B2 (en) 2015-12-04 2020-08-04 Raytheon Technologies Corporation Enhanced adhesion thermal barrier coating
US11306600B2 (en) 2015-12-04 2022-04-19 Raytheon Technologies Corporation Enhanced adhesion thermal barrier coating
CN111805715A (zh) * 2020-07-23 2020-10-23 王清冬 一种陶瓷喷釉机
CN111805715B (zh) * 2020-07-23 2021-09-03 内蒙古元史文化传媒有限公司 一种陶瓷喷釉机

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