Connect public, paid and private patent data with Google Patents Public Datasets

Method for controlling coating delamination caused when forming cooling holes through thermal barrier coatings

Info

Publication number
WO2016105327A1
WO2016105327A1 PCT/US2014/071784 US2014071784W WO2016105327A1 WO 2016105327 A1 WO2016105327 A1 WO 2016105327A1 US 2014071784 W US2014071784 W US 2014071784W WO 2016105327 A1 WO2016105327 A1 WO 2016105327A1
Authority
WO
Grant status
Application
Patent type
Prior art keywords
channels
layer
cooling
tbc
holes
Prior art date
Application number
PCT/US2014/071784
Other languages
French (fr)
Inventor
David G. Sansom
Ramesh Subramanian
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

Links

Classifications

    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01DNON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D5/00Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/12Blades
    • F01D5/28Selecting particular materials; Particular measures relating thereto; Measures against erosion or corrosion
    • F01D5/288Protective coatings for blades
    • CCHEMISTRY; METALLURGY
    • C23COATING METALLIC MATERIAL; COATING MATERIAL WITH METALLIC MATERIAL; CHEMICAL SURFACE TREATMENT; DIFFUSION TREATMENT OF METALLIC MATERIAL; COATING BY VACUUM EVAPORATION, BY SPUTTERING, BY ION IMPLANTATION OR BY CHEMICAL VAPOUR DEPOSITION, IN GENERAL; INHIBITING CORROSION OF METALLIC MATERIAL OR INCRUSTATION IN GENERAL
    • C23CCOATING METALLIC MATERIAL; COATING MATERIAL WITH METALLIC MATERIAL; SURFACE TREATMENT OF METALLIC MATERIAL BY DIFFUSION INTO THE SURFACE, BY CHEMICAL CONVERSION OR SUBSTITUTION; COATING BY VACUUM EVAPORATION, BY SPUTTERING, BY ION IMPLANTATION OR BY CHEMICAL VAPOUR DEPOSITION, IN GENERAL
    • C23C28/00Coating for obtaining at least two superposed coatings either by methods not provided for in a single one of groups C23C2/00 - C23C26/00 or by combinations of methods provided for in subclasses C23C and C25C or C25D
    • C23C28/30Coatings combining at least one metallic layer and at least one inorganic non-metallic layer
    • C23C28/32Coatings combining at least one metallic layer and at least one inorganic non-metallic layer including at least one pure metallic layer
    • C23C28/321Coatings combining at least one metallic layer and at least one inorganic non-metallic layer including at least one pure metallic layer with at least one metal alloy layer
    • C23C28/3215Coatings combining at least one metallic layer and at least one inorganic non-metallic layer including at least one pure metallic layer with at least one metal alloy layer at least one MCrAlX layer
    • CCHEMISTRY; METALLURGY
    • C23COATING METALLIC MATERIAL; COATING MATERIAL WITH METALLIC MATERIAL; CHEMICAL SURFACE TREATMENT; DIFFUSION TREATMENT OF METALLIC MATERIAL; COATING BY VACUUM EVAPORATION, BY SPUTTERING, BY ION IMPLANTATION OR BY CHEMICAL VAPOUR DEPOSITION, IN GENERAL; INHIBITING CORROSION OF METALLIC MATERIAL OR INCRUSTATION IN GENERAL
    • C23CCOATING METALLIC MATERIAL; COATING MATERIAL WITH METALLIC MATERIAL; SURFACE TREATMENT OF METALLIC MATERIAL BY DIFFUSION INTO THE SURFACE, BY CHEMICAL CONVERSION OR SUBSTITUTION; COATING BY VACUUM EVAPORATION, BY SPUTTERING, BY ION IMPLANTATION OR BY CHEMICAL VAPOUR DEPOSITION, IN GENERAL
    • C23C28/00Coating for obtaining at least two superposed coatings either by methods not provided for in a single one of groups C23C2/00 - C23C26/00 or by combinations of methods provided for in subclasses C23C and C25C or C25D
    • C23C28/30Coatings combining at least one metallic layer and at least one inorganic non-metallic layer
    • C23C28/34Coatings combining at least one metallic layer and at least one inorganic non-metallic layer including at least one inorganic non-metallic material layer, e.g. metal carbide, nitride, boride, silicide layer and their mixtures, enamels, phosphates and sulphates
    • C23C28/345Coatings combining at least one metallic layer and at least one inorganic non-metallic layer including at least one inorganic non-metallic material layer, e.g. metal carbide, nitride, boride, silicide layer and their mixtures, enamels, phosphates and sulphates with at least one oxide layer
    • C23C28/3455Coatings combining at least one metallic layer and at least one inorganic non-metallic layer including at least one inorganic non-metallic material layer, e.g. metal carbide, nitride, boride, silicide layer and their mixtures, enamels, phosphates and sulphates with at least one oxide layer with a refractory ceramic layer, e.g. refractory metal oxide, ZrO2, rare earth oxides or a thermal barrier system comprising at least one refractory oxide layer
    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23RGENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
    • F23R3/00Continuous combustion chambers using liquid or gaseous fuel
    • F23R3/002Wall structures
    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05DINDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2230/00Manufacture
    • F05D2230/90Coating; Surface treatment
    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05DINDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2300/00Materials; Properties thereof
    • F05D2300/60Properties or characteristics given to material by treatment or manufacturing
    • F05D2300/611Coating
    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23RGENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
    • F23R2900/00Special features of, or arrangements for continuous combustion chambers; Combustion processes therefor
    • F23R2900/00018Manufacturing combustion chamber liners or subparts

Abstract

A method for reducing defects on a component (105) having at least a substrate (116), a thermal barrier coating layer (112) over the substrate (116), and cooling holes (100) through the substrate and the thermal barrier coating layer (112) is provided. The method includes forming one or more channels (122) in the thermal barrier coating layer (112) about the perimeter of each cooling hole. The one or more channels (122) are effective to reduce crack propagation within the thermal barrier coating layer (112).

Description

METHOD FOR CONTROLLING COATING DELAM I NATION CAUSED WHEN FORMING COOLING HOLES THROUGH THERMAL BARRIER COATINGS

FIELD OF THE INVENTION

The invention relates generally to high temperature coatings, and in particular to methods for reducing damage to a thermal barrier coating on a substrate when forming cooling holes through the thermal barrier coating and substrate.

BACKGROUND OF THE INVENTION

A conventional combustion gas turbine engine includes a compressor, a combustor, and a turbine. The compressor compresses ambient air. The combustor combines the compressed air with a fuel and ignites the mixture creating combustion products defining a working gas. The working gas travels to the turbine, which includes a series of rows of stationary vanes and rotating blades. Each pair of rows of vanes and blades is referred to as a stage. Typically, there are four stages in a gas turbine. The rotating blades are coupled to a shaft and disc assembly. As the working gases expand through the turbine, the working gases cause the blades, and therefore the shaft and disc assembly, to rotate.

It is known that the efficiency of a combustion turbine engine will improve as the firing temperature of the combustion gas is increased. As the firing temperatures increase, the high temperature durability of the components of the turbine must increase correspondingly. Although nickel-based and cobalt-based superalloy materials are used for components in the hot gas flow path, such as combustor transition pieces and turbine rotating and stationary blades, even these superalloy materials are not capable of surviving long term operation at temperatures exceeding 1 ,400°C. In many applications, a superalloy substrate is coated with a thermal barrier coating (TBC) such as a ceramic insulating material in order to reduce the service temperature of the underlying substrate and to reduce the magnitude of the temperature transients to which the substrate is exposed.

In addition to a TBC, some conventional gas turbine components may employ cooling techniques such as a plurality of cooling holes that extend from an interior cooling passage in the component to an outer surface of the component to reduce the service temperature of the underlying substrate. The cooling holes may be shaped and sized such that the cooling fluid (e.g. a pressurized cooling fluid stream such as a compressed air stream) travels through each cooling hole. Typically, as shown in Prior Art FIG. 1 , the cooling holes 10 are formed through the thermal barrier coating (TBC) layer 12, a bond coat layer 14 (when present), and a substrate 16 at an angle less than 90 degrees through the layers.

The cooling holes 10 may be formed by laser drilling. In a number of cases, however, laser drilling through the TBC 12, bond coat 14, and substrate 16 to form the holes 10 can lead to spallation of the TBC 12 from the bond coat 14. In particular, spallation may occur when cracks 18 formed at the interface 20 of the TBC 12 and the bond coat 14 propagate and join. The cracking may be a result of one or more of thermal expansion, physical expulsion of material, and delamination effects caused by the laser melt. Further, cracking may be initiated or furthered during normal operation of the component. In any case, crack defects and loss of TBC material can lead to the bond coat layer 14 and substrate 16 being exposed to higher temperatures than desired, and may lead to premature degradation of the component.

BRIEF DESCRIPTION OF THE DRAWINGS

The invention is explained in the following description in view of the drawings that show:

FIG. 1 illustrates typical damage at an interface between a bond coat and a TBC layer when cooling holes are formed in the coating of a PRIOR ART component.

FIG. 2 illustrates a component comprising a plurality of channels adjacent cooling holes in accordance with an aspect of the present invention.

FIG. 3 illustrates a depth profile of channels formed in a TBC layer in accordance with an aspect of the present invention.

FIG. 4 illustrates channels having a circular shape about respective cooling holes in accordance with an aspect of the present invention.

FIG. 5 illustrates a channel having a semi-circular shape about a respective cooling hole in accordance with an aspect of the present invention. FIG. 6 illustrates a channel having a diamond shape about a respective cooling hole in accordance with an aspect of the present invention.

FIG. 7 illustrates a channel having straight lines about a respective cooling hole in accordance with an aspect of the present invention.

FIG. 8 illustrates a channel having curved lines about a respective cooling hole in accordance with an aspect of the present invention.

FIG. 9 illustrates a channel having a grid pattern about a cooling hole in accordance with an aspect of the present invention.

FIG. 10 illustrates grid patterns disposed about a plurality of cooling holes in accordance with an aspect of the present invention.

FIG. 1 1 is a graph showing a reduction in strain energy release rate as a result of the channels in accordance with an aspect of the present invention.

DETAILED DESCRIPTION OF THE INVENTION

The present inventors have found that forming channels into a depth of a thermal barrier coating (TBC) layer in a component comprising at least a substrate and a TBC layer thereon significantly reduces cracking associated with the formation of cooling holes through the layers. In certain embodiments, the channels may be disposed about at least a portion of the perimeter of the cooling holes to provide a controlled path for crack propagation to take place. In this way, instead of allowing cracks to propagate along an entire interface between the TBC layer and the bond coat or substrate, crack propagation may instead be directed (at least to an extent) toward the channels formed in the TBC layer, which significantly reduces the size and scale of any actual or potential spallation of the TBC layer. Further, the presence of the channels alone may reduce stress in the TBC layer to a degree similar to the reduction of a thickness of the TBC layer. The improved components (and processes for forming the same) having the channels formed therein may have higher temperature ratings and improved

performance profiles relative to components without the described channels.

Now referring to the figures, there is shown in FIG. 2 a component 105 comprising a substrate 1 16 and a TBC layer 1 12 disposed on the substrate 1 16. In the embodiment shown in FIG. 2, a bond coat layer 1 14 may be disposed on the substrate 1 16 between the substrate 1 16 and the TBC layer 1 12 to facilitate adherence of the TBC layer 1 12 to the substrate 16 and to reduce oxidation of the underlying substrate 1 16. In other embodiments, it is appreciated that the TBC layer 1 12 may be directly applied on the substrate 1 16 and no bond coat layer 1 14 is present. A plurality of cooling holes 100 are disposed through the TBC 1 12, bond coat layer 1 14 (when present), and the substrate 1 16.

To reduce spallation of the TBC layer 1 12 from the bond coat layer 1 14 or the substrate 1 16 and crack propagation within the TBC layer 1 12, the component 105 comprises a plurality of channels 122 disposed about at least a portion of a perimeter of at least some of the cooling holes 100 in a predetermined configuration. The channels 122 penetrate into a depth of the TBC layer 1 12, but not into a depth of the bond coat 1 14 (when present) or into a depth of the substrate 1 16. In certain embodiments, instead of allowing cracks to propagate along an entire interface 120 between the TBC layer 1 12 and the bond coat layer 1 14 (when present) or between the TBC layer 1 12 and the substrate 1 16, the channels 122 are effective to direct propagation of cracks 1 18 originating from the cooling holes 100 or originating from the interface 120 to respective channel(s) 122. It is also possible for the channels 122 to provide a controlled path for any crack propagation that may occur between adjacent channels 122. In this way, instead of allowing the crack propagation to travel along an entire length of the interface 120 from cooling hole 10 to cooling hole 10 as shown in FIG. 1 , for example, cracks 1 18 are instead provided with an outlet via the channels 122 to substantially reduce the extent of crack propagation.

In certain embodiments, any cracks 1 18 that form in the TBC layer 1 12 may propagate along the interface 120 only until it reaches a point close to the channel 122 where the channel 122 then acts as a path of least resistance. This results in the crack 1 18 turning upward toward the channel 122. The reduced crack propagation will in turn result in a substantially lesser likelihood of spallation, delamination, and thus loss of portions of the TBC layer 1 12.

In accordance with another aspect of the invention, the channels 122 are further effective to reduce crack propagation within the TBC layer 122 by reducing the driving force available for crack propagation. Referring to FIG. 1 1 , for example, the strain energy release rate is reduced by the presence of the channels 122 in the manner shown. In addition, the inventors have also found that the stress reduction provided by the channels 122 in certain embodiments may be generally equivalent to or

commensurate with a stress reduction provided by reducing a thickness of the TBC layer 1 12. In all, the channels 122 may reduce (such as by 50% relative to a TBC layer 1 12 without the channels 122) the driving force present to propagate a given crack further. Thus, even if the channels 122 do not fully cause deflection of the crack to the channels, the presence of the channels 122 alone reduce stresses in the TBC layer 1 12 and may reduce crack propagation in this manner.

The component 105 may be any structure that is subjected to an elevated temperature during operation. Thus, the component 105 may comprise a blade, a vane, a transition piece, or other component of a gas turbine, for example. The substrate 1 16 of the component 105 may be formed from any suitable material which would benefit from any of the processes described herein. In certain embodiments, the substrate 1 16 comprises a superalloy material. As noted above, the term "superalloy" is used herein as it is commonly used in the art to refer to a highly corrosion-resistant and oxidation-resistant alloy that exhibits excellent mechanical strength and resistance to creep even at high temperatures. Exemplary superalloys include, but are not limited to alloys sold under the trademarks and brand names Hastelloy, Inconel alloys (e.g., IN 738, IN 792, IN 939), Rene alloys (e.g. Rene N5, Rene 41 , Rene 80, Rene 108, Rene 142, Rene 220), Haynes alloys, Mar M, CM 247, CM 247 LC, C263, 71 8, X-750, ECY 768, 262, X45, PWA 1483 and CMSX (e.g. CMSX-4) single crystal alloys, GTD 1 1 1 , GTD 222, MGA 1400, MGA 2400, PSM 1 16, CMSX-8, CMSX-10, PWA 1484, IN 713C, Mar-M-200, PWA 1480, I N 100, IN 700, Udimet 600, Udimet 500, and titanium aluminide. Prior to the application of any materials thereto, the surface of the substrate 1 16 may be cleaned to remove contamination such as by aluminum oxide grit blasting.

As mentioned above, a bond coat layer 1 14 may be applied to a surface of the substrate 1 16 in order to improve the adhesion of a subsequently applied TBC layer 1 12 and to reduce the oxidation of the underlying substrate 1 16. Alternatively, the bond coat 1 14 may be omitted and the TBC layer 1 12 may be applied directly onto a surface of the substrate 1 16. An exemplary bond coat layer 1 14 comprises an MCrAlY material, where M denotes nickel, cobalt, iron, or mixtures thereof, Cr denotes chromium, Al denotes aluminum, and Y denotes yttrium. Another exemplary bond coat 1 14 comprises alumina. The bond coat 1 14 may be applied to the substrate 1 16 by any known process, such as sputtering, plasma spray, or electron beam physical vapor deposition.

The TBC layer 1 12 may comprise a yttria-stabilized zirconia (YSZ), which includes zirconium oxide (Zr02) with a predetermined concentration of yttrium oxide (Y2O3), pyrochlores, or other TBC material known in the art. The TBC layer 1 12 may be applied by an air plasma spray process, electron beam physical vapor deposition (EB- PVD) process, or any other suitable deposition technique, for example. An EB-PVD process typically provides the TBC layer 1 12, such as a YSZ coating, with a columnar microstructure having sub-micron sized gaps between adjacent columns of YSZ material as shown, for example, in U.S. Patent No. 5,562,998, the entirety of which is incorporated by reference herein. It is contemplated that the TBC layer 1 12 may have any desired thickness suitable for its intended application. In certain embodiments, the TBC layer 12 may have a thickness of from 1 -100 mils, and in a particular embodiment from 25-50 mils.

Referring again to FIG. 2, the cooling holes 100 may be defined within a portion of an outer wall of the component 105. For example, the cooling holes 100 may extend from an outer surface 126 of the substrate 1 16 through the TBC layer 1 12, through the bond coat layer 1 14 (when present), and through an inner surface 128 of the outer wall. The cooling holes 100 are typically in fluid communication with a flow path, such as an air channel 1 30, which continuously flows a pressurized fluid stream to the cooling holes 100 as shown. Advantageously, in operation, the cooling holes 100 form a fluid barrier between the component 105 and hot gases traveling through a main flowpath of the engine.

The cooling holes 100 may be formed by any suitable process known in the art such as by processes described in U.S. Patent Nos. 6,420,677; 6,339,208; 4,808,785; and 4,737,613, and U.S. Published Patent Application No. 2009/0074588, for example, the entirety of each which is hereby incorporated by reference. For example, the cooling holes 100 may be formed by directing an energy source at a predetermined point on the component for a predetermined period of time to form each cooling hole. The energy source or the component may then be advanced with respect to the other of the energy source or the component, and another hole may be formed in the same manner. Each of the cooling holes 100 may be formed so as to have substantially the same size, depth, angle relative to a surface (e.g. surface 126), and shape.

Alternatively, at least some of the cooling holes 1 00 may have different dimensions. It is appreciated that the depth, size, shape, and angle relative to a surface of the cooling holes 100 may be selected for the particular application as would be appreciated by the skilled artisan.

The channels 122 may be formed by any suitable process such as via any one of the processes described set forth in U.S. Patent No. 6,703,137, the entirety of which is hereby incorporated by reference herein. In one aspect, the channels 122 not only provide regions in which crack propagation may be confined, but also allow the TBC layer 1 12 to withstand a relatively large temperature gradient across its thickness without failure since the expansion and/or contraction of the TBC material can also be relieved by the channels 122.

The channels 122 may have a shape, depth, length, and width suitable for its intended application. In certain embodiments, the channels 122 do not extend all the way through a depth of the TBC layer 1 12 to the bond coat layer 1 14 (when present) or substrate 1 16 as shown in FIG. 2. In this way, exposure of the substrate 1 16 or the bond coat layer 1 14 to the environment of the component 105 is avoided. In an embodiment, the channels 122 have a depth of less than 75% of a total depth of the TBC layer 1 12, and in an embodiment, from 50-75 % of a total depth of the TBC layer 1 12. In addition, the channels 122 may be formed with any suitable angle relative to a surface of the TBC layer 1 12. In a particular embodiment, the channels 122 may be formed at an angle of from 25 degrees to 90 degrees relative to the top surface 126 of the TBC layer 1 12.

The selection of a particular channel 122 formation strategy, including the angle, dimensions, number, and locations of the channels 122, may vary from application to application, but should be selected to result in a level of stress within the thermal barrier coating 1 12 which is within allowable levels for the TBC layer 1 12 for the predetermined temperature environment.

The channels 122 may be formed in the TBC 1 12 utilizing a suitable energy sources such as a laser energy source. In certain embodiments, the laser source comprises a YAG laser having a wavelength of about 1.6 microns, for example. In this way, the YAG may serve as a finer cutting instrument than would a carbon dioxide laser which has a wavelength of about 10.1 microns, for example.

In operation, energy may be directed from the laser source toward the a top surface of the TBC layer 1 12 in order to heat the material in a localized area to a temperature sufficient to cause vaporization and removal of material to a desired depth to form the channels 122. It is appreciated that more than one interval (pass) may be necessary to form the channels 122 to the desired dimensions. The edges of the TBC material bounding the channels 122 may exhibit a small recast surface where material had been heated to just below the temperature necessary for vaporization.

The location of the channels 122 on the component 105, distance of the channels 122 from adjacent cooling holes 100 in the finished product, distance between adjacent channels 122, and the geometry of the channels 122 may be established by controlling parameters of the energy source. In an embodiment, the width of the channels 122 at the surface 130 of the TBC layer 1 12 is less than 50 microns, and in certain embodiments less than 25 microns. Wider, deeper, or narrower channel widths may also be selected for a component depending upon the sensitivity of the

aerodynamic design and the predicted thermal conditions. Preferably, the channels 122 are sized to provide a controlled path for crack propagation while having a minimal impact on aerodynamic efficiency.

The channels 122 may be disposed uniformly over all or a predetermined portion of the TBC layer 1 12 or may be disposed only in selected areas of the component 105. In addition, the channels 122 may further comprise relatively uniform shapes and dimensions on the same component or may comprise a mixture of channels having distinct sizes, patterns, and/or dimensions, for example.

Still further, the channels 122 are formed into a depth of the TBC layer 1 12 such that one or more channels 122 are provided about selected cooling holes 100 in the final product in a predetermined configuration. The predetermined configuration, dimensions thereof, and orientation thereof are without limitation. For exemplary purposes only, the predetermined configuration may comprise a member selected from the group consisting of a circle, a semi-circle, a diamond shape, a straight segment, a curved segment, and a grid pattern. It is thus understood that each predetermined configuration may comprise a single channel 122 (e.g., a single circle) or a plurality of channels 122 (e.g., a first circle inside a larger circle, or a grid). FIG. 4 illustrates a top view of a plurality of channels 122 disposed in the TBC layer 1 12 about respective cooling holes 100, each channel 122 formed in a predetermined configuration 130. In this embodiment, the predetermined configuration 130 comprises a single closed circle shape 132. In this embodiment, each closed circle shaped channel 132 is disposed about a respective cooling hole 100 and spaced apart therefrom by the distance (D).

In other embodiments, the predetermined configuration 130 may comprise one or more semi-circle shapes 134 (FIG. 5); one or more diamond shapes 136 (FIG. 6); one or more straight line segments 138 (FIG. 7); one or more curved segments 140 (FIG. 8); and/or one or more grid patterns 142 (FIG. 9). For ease of understanding, the configurations are shown about a single cooling hole 100 in FIGS. 5-9, but it is contemplated that the component 105 will comprise one or more channels 122 in a predetermined configuration 130 about at least a portion, if not all, of the cooling holes 100 formed in the component 105. It is further contemplated that the channels 122 described herein may intersect with the cooling holes 100 (e.g., FIG. 9) or may be spaced apart from the cooling holes 100 by a distance (e.g., distance (D) in FIG. 4).

FIG. 10 shows in further detail a plurality of channels 122 formed in the TBC layer 1 12 in a grid pattern 142 about each one of a plurality cooling holes 100. The channels 122 includes channels 144 traveling in a first direction and channels 146 which intersect the channels 144 traveling in the first direction and extend transverse to the first direction. Advantageously, the grid pattern 142 each provides a plurality of paths to which cracks within the TBC layer 1 12 may propagate and be confined. In this way, crack propagation within the TBC layer 1 12 can be controlled. It is appreciated that the grid pattern 142 may be oriented at any angle with respect to the cooling holes 100 in the final formed product. In certain embodiments, the plurality of channels 122 defining the grid pattern 142 may be oriented at an angle other than 90 degrees to a point on the perimeter of a respective cooling hole).

In accordance with another aspect, there is also provided a method for reducing defects on a component 100, including spallation of a TBC layer 1 12 on the component 100. The method comprises forming one or more channels 122 in the thermal barrier coating layer 1 12; and forming a cooling hole 100 extending through the thermal barrier coating layer 1 12 and the substrate 1 16. Upon formation of the one or more channels and the cooling hole 100, the one or more channels 122 are disposed about a perimeter of the cooling hole 100. The one or more channels 122 are effective to reduce crack propagation in the component 105 relative to the degree of crack propagation in a component without the channels. For example, the one or more channels 1 12 may reduce crack propagation by directing crack propagation from the cooling hole 100 to the one or more channels 122 upon an occurrence of such crack propagation. In addition, the channels 122 may be effective to reduce crack propagation within the TBC layer 122 by reducing the driving force available for crack propagation. In an

embodiment, a plurality of cooling holes are formed and one or more channels are disposed about at least a plurality of the cooling holes.

It is contemplated that the forming of the plurality of cooling holes 100 may be performed following the forming of the plurality of channels 122. In this way, the channels 1 12 are already present and in their desired position to optimally limit crack propagation which occurs during the formation of any cooling holes 100. In another embodiment, the forming of the plurality of cooling holes 100 is done prior to the forming of the plurality of channels 122. In this way, the methods described herein can be applied to existing components having cooling holes already formed therein to reduce any further spallation of the TBC layer 1 12 of the component 105.

While various embodiments of the present invention have been shown and described herein, it will be obvious that such embodiments are provided by way of example only. Numerous variations, changes and substitutions may be made without departing from the invention herein. Accordingly, it is intended that the invention be limited only by the spirit and scope of the appended claims.

Claims

The invention claimed is: 1 . A method for reducing defects on a component including a substrate, a thermal barrier coating layer over the substrate, and cooling holes through the substrate and the thermal barrier coating layer, the method comprising:
forming one or more channels in the thermal barrier coating layer about the cooling holes, the one or more channels effective to reduce crack propagation within the component.
2. The method of claim 1 , wherein the one or more channels are formed in a predetermined configuration selected from the group consisting of one or more circles, one or more semi-circles, one or more diamond shapes, one or more straight segments, one or more curved segments, and one or more grid patterns.
3. The method in any of claims 1 to 2, wherein a plurality of the channels are formed in a grid pattern.
4. The method in any of claims 1 to 3, wherein the one or more channels are formed by a YAG laser source.
5. The method in any of claims 1 to 4, wherein the one or more channels have a depth that is 75% or less than an entire depth of the thermal barrier coating layer.
6. The method of claim 5, wherein the one or more channels have a depth from 50-75 % of an entire depth of the thermal barrier coating layer.
7. The method in any of claims 1 to 6, wherein the one or more channels are formed at an angle of from 25 to 90 degrees relative to a top surface of the thermal barrier coating layer.
8. A method for reducing defects on a component including a substrate and a thermal barrier coating layer thereon, the method comprising:
forming one or more channels in the thermal barrier coating layer; and
forming a cooling hole extending through the thermal barrier coating layer and the substrate; and
wherein, upon formation of the one or more channels and the cooling hole, the one or more channels are disposed about a perimeter of the cooling hole and the one or more channels are effective to reduce crack propagation within the component.
9. The method of claim 8, wherein the forming of the cooling hole is done following the forming of the one or more channels.
10. The method of claim 8, wherein the forming of the cooling hole is done prior to the forming of the one or more channels.
1 1 . The method in any of claims 8 to 10, wherein the one or more channels are formed in a predetermined configuration selected from the group consisting of one or more circles, one or more semi-circles, one or more diamond shapes, one or more straight segments, one or more curved segments, and one or more grid patterns.
12. The method of claim 1 1 , wherein a plurality of the channels are formed in a grid pattern.
13. The method in any of claims 8 to 12, wherein the one or more channels are formed by a YAG laser source.
14. The method in any of claims 8 to 13, wherein the channel has a depth that is 75% or less than an entire depth of the thermal barrier coating layer.
15. The method in any of claims 8 to 14, wherein the component comprises a plurality of cooling holes, and wherein one or more channels are disposed about each of the plurality of cooling holes.
16. A component comprising:
a substrate;
a thermal barrier coating layer disposed on the substrate;
a cooling hole extending through the thermal barrier coating and the substrate; one or more channels formed in the thermal barrier coating layer about the cooling holes, the one or more channels effective to reduce crack propagation within the component.
17. The component of claim 16, wherein the one or more channels are formed in a predetermined configuration selected from the group consisting of one or more circles, one or more semi-circles, one or more diamond shapes, one or more straight segments, one or more curved segments, and one or more grid patterns.
18. The component in any of claims 16 to 17, wherein a plurality of the channels are formed in a grid pattern about the cooling hole.
19. The component in any of claims 16 to 18, wherein the component comprises a gas turbine component selected from the group consisting of a blade, a vane, and a transition piece.
20. The component in any of claims 16 to 19, further comprising a bond coat layer disposed between the thermal barrier coating layer and the substrate.
21 . The gas turbine in any of claims 16 to 20, wherein the one or more channels have a depth of from 50-75 % of a total depth of the thermal barrier coating layer.
22. The gas turbine in any of claims 16 to 21 , wherein the one or more channels are formed at an angle of from 25 to 90 degrees relative to a top surface of the thermal barrier coating layer.
PCT/US2014/071784 2014-12-22 2014-12-22 Method for controlling coating delamination caused when forming cooling holes through thermal barrier coatings WO2016105327A1 (en)

Priority Applications (1)

Application Number Priority Date Filing Date Title
PCT/US2014/071784 WO2016105327A1 (en) 2014-12-22 2014-12-22 Method for controlling coating delamination caused when forming cooling holes through thermal barrier coatings

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
PCT/US2014/071784 WO2016105327A1 (en) 2014-12-22 2014-12-22 Method for controlling coating delamination caused when forming cooling holes through thermal barrier coatings

Publications (1)

Publication Number Publication Date
WO2016105327A1 true true WO2016105327A1 (en) 2016-06-30

Family

ID=52394352

Family Applications (1)

Application Number Title Priority Date Filing Date
PCT/US2014/071784 WO2016105327A1 (en) 2014-12-22 2014-12-22 Method for controlling coating delamination caused when forming cooling holes through thermal barrier coatings

Country Status (1)

Country Link
WO (1) WO2016105327A1 (en)

Citations (12)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US4737613A (en) 1987-08-24 1988-04-12 United Technologies Corporation Laser machining method
US4808785A (en) 1986-11-13 1989-02-28 Chromalloy Gas Turbine Corporation Method and apparatus for making diffused cooling holes in an airfoil
US5562998A (en) 1994-11-18 1996-10-08 Alliedsignal Inc. Durable thermal barrier coating
US6339208B1 (en) 2000-01-19 2002-01-15 General Electric Company Method of forming cooling holes
US6420677B1 (en) 2000-12-20 2002-07-16 Chromalloy Gas Turbine Corporation Laser machining cooling holes in gas turbine components
US6703137B2 (en) 2001-08-02 2004-03-09 Siemens Westinghouse Power Corporation Segmented thermal barrier coating and method of manufacturing the same
US20090017260A1 (en) * 2001-08-02 2009-01-15 Kulkarni Anand A Segmented thermal barrier coating
US20090074588A1 (en) 2007-09-19 2009-03-19 Siemens Power Generation, Inc. Airfoil with cooling hole having a flared section
EP2275645A2 (en) * 2009-07-17 2011-01-19 Rolls-Royce Corporation Gas turbine component comprising stress mitigating features
US20130209233A1 (en) * 2012-02-15 2013-08-15 United Technologies Corporation Cooling hole with enhanced flow attachment
EP2733236A1 (en) * 2012-11-16 2014-05-21 Siemens Aktiengesellschaft Two-layer ceramic coating system having an outer porous layer and depressions therein
EP2733310A1 (en) * 2012-11-16 2014-05-21 Siemens Aktiengesellschaft Modified surface around a hole

Patent Citations (12)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US4808785A (en) 1986-11-13 1989-02-28 Chromalloy Gas Turbine Corporation Method and apparatus for making diffused cooling holes in an airfoil
US4737613A (en) 1987-08-24 1988-04-12 United Technologies Corporation Laser machining method
US5562998A (en) 1994-11-18 1996-10-08 Alliedsignal Inc. Durable thermal barrier coating
US6339208B1 (en) 2000-01-19 2002-01-15 General Electric Company Method of forming cooling holes
US6420677B1 (en) 2000-12-20 2002-07-16 Chromalloy Gas Turbine Corporation Laser machining cooling holes in gas turbine components
US6703137B2 (en) 2001-08-02 2004-03-09 Siemens Westinghouse Power Corporation Segmented thermal barrier coating and method of manufacturing the same
US20090017260A1 (en) * 2001-08-02 2009-01-15 Kulkarni Anand A Segmented thermal barrier coating
US20090074588A1 (en) 2007-09-19 2009-03-19 Siemens Power Generation, Inc. Airfoil with cooling hole having a flared section
EP2275645A2 (en) * 2009-07-17 2011-01-19 Rolls-Royce Corporation Gas turbine component comprising stress mitigating features
US20130209233A1 (en) * 2012-02-15 2013-08-15 United Technologies Corporation Cooling hole with enhanced flow attachment
EP2733236A1 (en) * 2012-11-16 2014-05-21 Siemens Aktiengesellschaft Two-layer ceramic coating system having an outer porous layer and depressions therein
EP2733310A1 (en) * 2012-11-16 2014-05-21 Siemens Aktiengesellschaft Modified surface around a hole

Similar Documents

Publication Publication Date Title
DeMasi-Marcin et al. Protective coatings in the gas turbine engine
US5735044A (en) Laser shock peening for gas turbine engine weld repair
US6461107B1 (en) Turbine blade tip having thermal barrier coating-formed micro cooling channels
US6933061B2 (en) Thermal barrier coating protected by thermally glazed layer and method for preparing same
US5732467A (en) Method of repairing directionally solidified and single crystal alloy parts
US6528118B2 (en) Process for creating structured porosity in thermal barrier coating
EP1283278A2 (en) Segmented thermal barrier coating and method of manufacturing the same
EP2275645A2 (en) Gas turbine component comprising stress mitigating features
US20120148769A1 (en) Method of fabricating a component using a two-layer structural coating
US20120051941A1 (en) Components with conformal curved film holes and methods of manufacture
US20080028605A1 (en) Weld repair of metallic components
US20090324841A1 (en) Method of restoring near-wall cooled turbine components
US20060016191A1 (en) Combined effusion and thick TBC cooling method
US20130232749A1 (en) Advanced pass progression for build-up welding
US6933066B2 (en) Thermal barrier coating protected by tantalum oxide and method for preparing same
US7509735B2 (en) In-frame repairing system of gas turbine components
WO2007112783A1 (en) Layered thermal barrier coating with a high porosity, and a component
US20120111545A1 (en) Components with re-entrant shaped cooling channels and methods of manufacture
US20030021905A1 (en) Method for cooling engine components using multi-layer barrier coating
US8533949B2 (en) Methods of manufacture for components with cooling channels
US20110038710A1 (en) Application of Dense Vertically Cracked and Porous Thermal Barrier Coating to a Gas Turbine Component
US20120124832A1 (en) Turbine components with cooling features and methods of manufacturing the same
EP1685923A1 (en) Repair and reclassification of superalloy components
US6893750B2 (en) Thermal barrier coating protected by alumina and method for preparing same
US20130101761A1 (en) Components with laser cladding and methods of manufacture

Legal Events

Date Code Title Description
121 Ep: the epo has been informed by wipo that ep was designated in this application

Ref document number: 14828619

Country of ref document: EP

Kind code of ref document: A1

NENP Non-entry into the national phase in:

Ref country code: DE