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`Samsung Electronics Co., Ltd. v. Demaray LLC
`Samsung Electronic's Exhibit 1038
`Exhibit 1038, Page 1
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`U.S. Patent
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`Sep. 30, 2003
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`US 6,627,323 B2
`
` 2
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`FIG,
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`Ex. 1038, Page 2
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`Ex. 1038, Page 2
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`US 6,627,323 B2
`
`
`1
`THERMAL BARRIER COATING RESISTANT
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`TO DEPOSITS AND COATING METHOD
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`THEREFOR
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`2
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`with compounds found within a gas turbine engine during its
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`operation. Notable contaminants include such oxides as
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`calcia, magnesia, alumina and silica, which whenpresent
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`together at elevated temperatures form a compoundreferred
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`to herein as CMAS. CMAShasa relatively low melting
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`eutectic (about 1190° C.)
`that when molten is able to
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`infiltrate to the cooler subsurface regions of a TBC, whereit
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`resolidifies. During thermal cycling,
`the CTE mismatch
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`between CMASand the TBC promotes spallation, particu-
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`larly TBC deposited by PVD and APSdueto the ability of
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`the molten CMASto penetrate their columnar and porous
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`grain structures, respectively. Another detriment of CMASis
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`that the bond coat and substrate underlying the TBC are
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`susceptible to corrosion attack by alkali deposits associated
`with the infiltration of CMAS.
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`Various studies have been performed to find coating
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`materials that are resistant to infiltration by CMAS. Notable
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`examples are U.S. Pat. Nos. 5,660,885, 5,871,820 and
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`5,914,189 to Hasz et al., which disclose three types of
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`coatings to protect a TBC from CMAS-related damage.
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`These protective coatings are classified as being
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`impermeable, sacrificial or non-wetting to CMAS. Imper-
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`meable coatings are defined as inhibiting infiltration of
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`molten CMAS,and includesilica, tantala, scandia, alumina,
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`hafnia, zirconia, calcium zirconate, spinels, carbides,
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`nitrides, silicides, and noble metals such as platinum. Sac-
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`rificial coatings are said to react with CMASto increase the
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`melting temperature or the viscosity of CMAS,
`thereby
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`inhibiting infiltration. Suitable sacrificial coating materials
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`include silica, scandia, alumina, calcium zirconate, spinels,
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`magnesia, calcia and chromia. As its name implies, a non-
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`wetting coating is non-wetting to molten CMAS, with
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`suitable materials including silica, hafnia, zirconia, beryl-
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`lium oxide, lanthana, carbides, nitrides, silicides, and noble
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`metals such as platinum. According to the Hasz et al.
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`patents, an impermeable coating or a sacrificial coating is
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`deposited directly on the TBC, and may be followed by a
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`layer of impermeable coating (if a sacrificial coating was
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`deposited first), sacrificial coating (if the impermeable coat-
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`ing was depositedfirst), or non-wetting coating. If used, the
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`non-wetting coating is the outermost coating of the protec-
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`tive coating system.
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`While the coating systems disclosed by Hasz et al. are
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`effective in protecting a TBC from damageresulting from
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`CMASinfiltration, further improvements would be desir-
`able.
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`Ex. 1038, Page 3
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`CROSS REFERENCE TO RELATED
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`
`APPLICATIONS
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`Not applicable.
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`STATEMENT REGARDING FEDERALLY
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`SPONSORED RESEARCH
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`10
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`Not applicable.
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`15
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`20
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`25
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`30
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`35
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`40
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`45
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`50
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`55
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`65
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`BACKGROUND OF THE INVENTION
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`1. Field of the Invention
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`This invention generally relates to coatings for compo-
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`nents exposed to high temperatures, such as the hostile
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`thermal environment of a gas turbine engine. More
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`particularly, this invention is directed to a protective coating
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`system for a thermal barrier coating on a gas turbine engine
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`component, in which the protective coating system is resis-
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`tant to infiltration by contaminants present in the operating
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`environmentof a gas turbine engine.
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`2. Description of the Related Art
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`Hot section components of gas turbine enginesare often
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`protected by a thermal barrier coating (TBC), which reduces
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`the temperature of the underlying component substrate and
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`thereby prolongsthe service life of the component. Ceramic
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`materials and particularly yttria-stabilized zirconia (YSZ)
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`are widely used as TBC materials because of their high
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`temperature capability, low thermal conductivity, and rela-
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`tive ease of deposition by plasma spraying, flame spraying
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`and physical vapor deposition (PVD) techniques. Air plasma
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`spraying (APS) has the advantagesof relatively low equip-
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`ment costs and ease of application and masking, while
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`TBC’s employed in the highest temperature regions of gas
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`turbine engines are often deposited by PVD, particularly
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`electron-beam PVD (EBPVD), whichyieldsa strain-tolerant
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`columnargrain structure. Similar columnar microstructures
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`can be produced using other atomic and molecular vapor
`processes.
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`To be effective, a TBC must strongly adhere to the
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`component and remain adherent throughout many heating
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`and cooling cycles. The latter requirement is particularly
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`demanding due to the different coefficients of thermal
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`expansion (CTE) between ceramic materials and the sub-
`BRIEF SUMMARY OF THE INVENTION
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`strates they protect, which are typically superalloys, though
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`ceramic matrix composite (CMC) materials are also used.
`invention generally provides a protective
`The present
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`coating system and method for protecting a thermal barrier
`An oxidation-resistant bond coat is often employed to pro-
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`mote adhesion and extend the service life of a TBC, as well
`coating (TBC) on a componentused in a high-temperature
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`environment, such as the hot section of a gas turbine engine.
`as protect the underlying substrate from damage by oxida-
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`The invention is particularly directed to a protective coating
`tion and hotcorrosion attack. Bond coats used on superalloy
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`system that significantly reduces if not prevents the infiltra-
`substrates are typically in the form of an overlay coating
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`such as MCrAIX (where M isiron, cobalt and/or nickel, and
`tion of CMASinto the underlying TBC.
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`X is yttrium or another rare earth element), or a diffusion
`The protective coating system of this invention comprises
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`aluminide coating. During the deposition of the ceramic
`inner and outer alumina layers and a platinum-group metal
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`TBC and subsequent exposures to high temperatures, such
`layer. The inner alumina layer is deposited on the thermal
`60
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`as during engine operation, these bond coats formatightly
`barrier coating, the platinum-group metal layer is deposited
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`adherent alumina (Al,O,) layer or scale that adheres the
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`on the inner alumina layer, and the outer alumina layer is
`TBC to the bond coat.
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`deposited on the platinum-group metal layer, so that the
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`The service life of a TBC system is typically limited by
`platinum-group metal layer is encased between the inner and
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`a spallation event driven by bond coat oxidation and the
`outer alumina layers. The outer aluminalayer is intended as
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`resulting thermal fatigue. In addition to the CTE mismatch
`a sacrificial layer that reacts with molten CMAS, forming a
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`between a ceramic TBC and a metallic substrate, spallation
`compound with a melting temperature that is significantly
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`can be promoted as a result of the TBC being contaminated higher than CMAS. Asaresult, the reaction product of the
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`Ex. 1038, Page 3
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`US 6,627,323 B2
`
`10
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`15
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`3
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`outer alumina layer and CMASresolidifies before it can
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`infiltrate the TBC. The platinum-group metal
`layer
`is
`believedto serve as a barrierto infiltration of CMASinto the
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`inner alumina layer and, therefore, the TBC. Notably, the
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`inner alumina layer beneath the platinum-group metal layer
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`appears to enhance the ability of the platinum-group metal
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`layer to prevent infiltration of CMAS. In other words, the
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`platinum-group metal layer is better able to perform as a
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`barrier to CMASinfiltration if it is deposited on an alumina
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`layer than if it were deposited directly on the TBC.
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`In view of the above, the protective coating system ofthis
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`invention is able to increase the temperature capability of a
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`TBC by reducing the vulnerability of the TBC to spallation
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`and the underlying substrate to corrosion from CMAS
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`contamination. The layers of the protective coating system
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`can be preferentially deposited on limited surface areas of a
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`component more susceptible to CMAS contamination. In
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`this manner, the additional weight and cost incurred by the
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`protective coating system can be minimized. Finally,
`the
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`protective coating system of this invention can be applied
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`during the process of rejuvenating a TBC on a component
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`returned from field service, thereby further extendingthe life
`of a TBC.
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`Other objects and advantages of this invention will be
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`better appreciated from the following detailed description.
`BRIEF DESCRIPTION OF THE DRAWINGS
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`FIG. 1 is a perspective view of a high pressure turbine
`blade.
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`FIG. 2 is a cross-sectional view of the blade of FIG. 1
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`along line 2—2, and showsa protective coating overlaying
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`a thermal barrier coating in accordance with this invention.
`DETAILED DESCRIPTION OF THE
`
`
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`INVENTION
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`4
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`such as EBPVD. The invention is also applicable to nonco-
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`lumnar TBC deposited by such methods as plasmaspraying,
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`including air plasma spraying (APS). A TBC ofthis type is
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`in the form of molten “splats,” resulting in a microstructure
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`characterized by irregular flattened grains and a degree of
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`inhomogeneity and porosity.
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`As with prior art TBC’s, the TBC 26 of this invention is
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`intended to be deposited to a thickness that is sufficient to
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`provide the required thermal protection for the underlying
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`substrate 22 and blade 10. A suitable thickness is generally
`on the order of about 75 to about 300 micrometers. A
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`preferred material for the TBC 26 is an yttria-stabilized
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`zirconia (YSZ), a preferred composition being about 3 to
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`about 8 weight percent yttria, though other ceramic mate-
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`rials could be used, such as nonstabilized zirconia, or
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`zircoma partially or fully stabilized by magnesia, ceria,
`scandia or other oxides.
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`invention is the
`to the present
`Of particular interest
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`susceptibility of TBC materials, including YSZ,to attack by
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`CMAS. As discussed previously, CMASis a relatively low
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`melting eutectic that when molten is able to infiltrate colum-
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`nar and porous TBC materials, and subsequently resolidify
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`to promote spallation during thermal cycling. To address this
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`concern, the TBC 26 in FIG. 2 is shownas being overcoated
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`by a protective coating system 30 of this invention. As the
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`outermost
`layer on the blade 10,
`the protective coating
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`system 30 serves as a barrier to CMASinfiltration of the
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`underlying TBC 26. The protective coating system 30 is
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`shown in FIG. 2 as comprising four discrete layers 32, 34,
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`36 and 38. The innermost layer 32 and the third layer 36 of
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`the coating system 30 are formed of alumina (Al,O;). The
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`layer 34 between the aluminalayers 32 and 36 is formed of
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`a platinum-group metal, which includes platinum,
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`ruthenium, rhodium, palladium, osmium and iridium. The
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`outermost layer 38 is an optional member of the coating
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`system 30, and is intended to provide a nonstick surface to
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`which CMASwill not readily wet and bond. A particularly
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`suitable material for the outermost layer 38 is believed to be
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`tantala, though it is foreseeable that other materials with
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`similar nonstick properties could be used. A suitable thick-
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`ness for the nonstick layer 38 is about 0.5 to about 5
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`micrometers, more preferably about 0.5 to about 2 microme-
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`As represented in FIG. 2, the alumina layers 32 and 36
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`have dense microstructures as a result of being deposited by
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`PVD, chemical vapor deposition (CVD) or another suitable
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`technique known in the art. The function of the inner and
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`outer alumina layers 32 and 36 is to serve as sacrificial
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`layers, reacting with molten CMASthatinfiltrates the pro-
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`tective coating system 30 to form one or more refractory
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`phases with higher melting temperatures than CMAS. In
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`effect, the alumina content of CMASis increased above the
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`eutectic point, yielding a modified CMAS with a higher
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`melting and/or crystallization temperature. As a result, the
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`reaction productof the inner and outer alumina layers 32 and
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`36 and CMAStendsto resolidify before infiltrating the TBC
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`26. Asuitable thickness for the outer alumina layer 36 is on
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`the order of about 0.5 to about 5 micrometers, more pref-
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`erably about 0.5 to about 2 micrometers, while a suitable
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`thickness for the inner alumina layer 32 is believed to be
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`about 0.5 to about 50 micrometers, more preferably about 5
`to about 10 micrometers.
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`The platinum-group metal layer 34 is believed to serve as
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`a barrier to infiltration of CMASinto the inner aluminalayer
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`32, thus enhancing the ability of the inner aluminalayer 32
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`to react with CMAS. A suitable method for depositing the
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`metal layer 34 is again a CVD or PVD technique such as
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`Ex. 1038, Page 4
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`20
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`25
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`30
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`35
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`The present invention will be described in reference to a
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`high pressure turbine blade 10 shown in FIG. 1, though the
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`invention is generally applicable to any component that
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`operates within a thermally and chemically hostile environ-
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`ment. The blade 10 generally includes an airfoil 12 against
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`which hot combustion gasesare directed during operation of
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`the gas turbine engine, and whose surfaces are therefore
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`subjected to severe attack by oxidation, hot corrosion and
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`erosion. The airfoil 12 is anchored to a turbine disk (not
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`shown) with a dovetail 14 formed ona rootsection 16 ofthe
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`blade 10. Cooling holes 18 are present in the airfoil 12
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`through which bleed air is forced to transfer heat from the
`blade 10.
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`The surface of the airfoil 12 is protected by a TBC system
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`20, represented in FIG. 2 as including a metallic bond coat
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`24 that overlies the surface of a substrate 22, the latter of
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`which may be a superalloy and typically the base material of
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`the blade 10. As widely practiced with TBC systems for
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`components of gas turbine engines,
`the bond coat 24 is
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`preferably an aluminum-rich composition, such as an over-
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`lay coating of an MCrAIXalloy or a diffusion coating such
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`as a diffusion aluminide or a diffusion platinum aluminide,
`all of which are knownin the art. Aluminum-rich bond coats
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`develop an aluminum oxide (alumina) scale 28, which is
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`grown by oxidation of the bond coat 24. The alumina scale
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`28 chemically bonds a TBC 26, formed of a thermal-
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`insulating material, to the bond coat 24 and substrate 22. The
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`TBC 26 of FIG. 2 is represented as having a strain-tolerant
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`microstructure of columnargrains. As knownin theart, such
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`columnar microstructures can be achieved by depositing the
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`TBC 26 using a physical vapor deposition (PVD) technique,
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`40
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`Ex. 1038, Page 4
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`US 6,627,323 B2
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`5
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`sputtering. The platinum-group metal layer 34 is preferably
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`entirely covered by the outer alumina layer 36, such that
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`platinum-group metal is not present at the external surface of
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`the coating system 30. With this arrangement, the outer
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`aluminalayer 36 serves to protect the platinum-group metal
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`layer 34 from degradation. Importantly, the presence of the
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`inner alumina layer 32 beneath the platinum-group metal
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`layer 34 appears to enhancethe ability of the platinum-group
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`metal layer 34 to prevent infiltration of CMAS. In other
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`words, improved resistant to CMASinfiltration appears to
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`be obtained if the platinum-group metal layer 34 is encased
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`between the alumina layers 32 and 34, in comparison to a
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`coating system in which the platinum-group metal layer is
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`directly deposited on a TBC oris the outermost layer of the
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`coating system. In its role as a barrier, a suitable thickness
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`for the platinum-group metal layer 34 is believed to be about
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`0.1 to about 2 micrometers, more preferably about 0.1 to
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`about 0.5 micrometers. To promote the adhesion of the
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`coating system 30, the surface of the TBC 26 is preferably
`20
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`polished prior to deposition of the inner aluminalayer 32.A
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`suitable surface finish is about 30 micro-inches (about 0.75
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`micrometers) Raor less.
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`There are various opportunities for making use of the
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`benefits of the protective coating system 30 ofthis invention.
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`For example, the coating system 30 can be applied to newly
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`manufactured components that have not been exposed to
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`service. Alternatively, the coating system 30 can be applied
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`to a componentthat has seen service, and whose TBC must
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`be cleaned and rejuvenated before being returned to the
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`field. In the latter case, applying the coating system 30 to the
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`TBC can significantly extend the life of the component
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`beyondthat otherwise possible if the TBC was not protected
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`by the coating system 30. In a preferred embodiment, the
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`coating system 30 is deposited only on those surfaces of a
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`componentthat are particularly susceptible to damage from
`CMASinfiltration. In the case of the blade 10 shownin FIG.
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`1, of particular interest
`is often the concave (pressure)
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`surface 40 of the airfoil 12, which is can be significantly
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`more susceptible to attack than the convex (suction) surface
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`42 as a result of aerodynamic considerations. According to
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`the invention,the layers 32, 34, 36 and optional layer 38 of
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`the coating system 30 can be selectively deposited on the
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`concave surface 40 of the airfoil 12, thus minimizing the
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`additional weight and cost of the coating system 30. For this
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`purpose, preferred deposition techniques include sputtering
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`and directed PVD. Multiple blades can be simultaneously
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`coated by positioning their convex surfaces back-to-back, so
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`that their convex surfaces effectively mask each other and
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`their concave surfaces face outward for coating. Deposition
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`by sputtering or directed PVD can then be performed to
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`deposit the coating system 30 essentially exclusively on the
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`exposed concave blade surfaces. While the concave surface
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`40 of the airfoil 12 may be of particular interest, circum-
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`stances may exist where other surface areas of the blade 10
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`are of concern, such as the leading edge of the airfoil 12 or
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`the region of the convex surface of the airfoil 12 near the
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`leading edge.
`While the invention has been described in terms of a
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`preferred embodiment,it is apparent that other forms could
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`be adopted by one skilled in the art, such as by substituting
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`other TBC,bond coat and substrate materials, or by utilizing
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`other methods to deposit and process the protective coating
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`system. Accordingly,
`the scope of the invention is to be
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`limited only by the following claims.
`Whatis claimedis:
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`1. A component having a thermal barrier coating on a
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`surface thereof,
`the component comprising a protective
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`10
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`15
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`30
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`35
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`40
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`50
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`65
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`6
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`coating system overlying the thermal barrier coating, the
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`protective coating system comprising inner and outer alu-
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`mina layers and a platinum-group metal
`layer encased
`therebetween.
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`2. Acomponentaccording to claim 1, wherein the thermal
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`barrier coating is yttria-stabilized zirconia.
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`3. A component according to claim 1, wherein the pro-
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`tective coating system consists of the inner and outer alu-
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`mina layers and the platinum-group metallayer.
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`4. A component according to claim 1, wherein the
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`platinum-group metal layer consists essentially of platinum.
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`5. A component according to claim 1, wherein the com-
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`ponentis an airfoil component of a gas turbine engine.
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`6. A component according to claim 5, wherein the com-
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`ponenthas a concavesurface, a convex surface and a leading
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`edge therebetween, and the protective coating system over-
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`lies only one of the concave surface, the convex surface or
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`the leading edge.
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`7. Acomponent according to claim 1, wherein the inner
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`alumina layer has a thickness of about 0.5 to about 50
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`micrometers, the platinum-group metal layer has a thickness
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`of about 0.1 to about 2 micrometers, and the outer alumina
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`layer has a thickness of about 0.5 to about 5 micrometers.
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`8. A component according to claim 1, wherein the pro-
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`tective coating system further comprises a layer of tantala
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`overlying the outer aluminalayer.
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`9. Acomponent according to claim 8, wherein the tantala
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`layer has a thickness of about 0.5 to about 5 micrometers.
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`10. A gas turbine engine component having a thermal
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`barrier coating of yttria-stabilized zirconia, the component
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`comprising an outer protective coating system overlying the
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`thermal barrier coating, the protective coating system com-
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`prising a platinum-group metal layer encased between inner
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`and outer alumina layers having columnargrain structures,
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`such that platinum-group metal is not present at an external
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`surface of the component defined by the protective coating
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`system.
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`11. A component according to claim 10, wherein the
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`protective coating system consists of the inner and outer
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`alumina layers and the platinum-group metal layer, and the
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`outer alumina layer defines the external surface of the
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`component.
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`12. A component according to claim 10, wherein the
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`platinum-group metal layer consists essentially of platinum.
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`13. A component according to claim 10, wherein the
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`component
`is an airfoil component having a concave
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`surface, a convex surface and a leading edge therebetween,
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`and the protective coating system overlies only one of the
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`concave surface, the convex surface or the leading edge.
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`14. A componentaccording to claim 10, wherein the inner
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`alumina layer has a thickness of about 5 to about 10
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`micrometers, the platinum-group metal layer has a thickness
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`of about 0.1 to about 0.5 micrometers, and the outer alumina
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`layer has a thickness of about 0.5 to about 2 micrometers.
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`15. A component according to claim 10, wherein the
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`protective coating system further comprises a layer of tan-
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`tala overlying the outer aluminalayer, and the tantala layer
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`defines the external surface of the component.
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`16. A component according to claim 15, wherein the
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`tantala layer has a thickness of about 0.5 to about 2
`micrometers.
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`17. Acomponentaccording to claim 10, wherein CMAS
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`
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`hasinfiltrated the columnargrains of the outer aluminalayer,
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`
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`the platinum-group metal layer being a barrier to infiltration
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`
`
`
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`
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`of the CMASinto the inner aluminalayer.
`
`
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`
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`18. A method of protecting a thermal barrier coating on a
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`
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`surface of a component, the method comprising the step of
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`Ex. 1038, Page 5
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`Ex. 1038, Page 5
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`7
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`depositing a protective coating system on the thermalbarrier
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`coating, the protective coating system comprising an inner
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`alumina layer deposited on the thermal barrier coating, a
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`platinum-group metal layer deposited on the inner alumina
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`layer, and an outer aluminalayer deposited on the platinum-
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`group metal layer so that the platinum-group metal layer is
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`encased between the inner and outer aluminalayers.
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`19. A method according to claim 18, wherein the thermal
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`barrier coating is yttria-stabilized zirconia.
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`20. A method according to claim 18, wherein the protec-
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`tive coating system consists of the inner and outer alumina
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`layers and the platinum-group metal layer.
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`21. A method according to claim 18, wherein the
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`platinum-group metal layer consists essentially of platinum.
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`22. A method according to claim 18, wherein the com-
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`ponentis an airfoil component of a gas turbine engine.
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`23. A method according to claim 22, wherein the com-
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`ponenthas a concavesurface, a convex surface and a leading
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`edge therebetween, and the protective coating system is
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`selectively deposited on only one of the concave surface, the
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`convex surface or the leading edge.
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`24. A method according to claim 23, wherein each layer
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`of the protective coating system is deposited by sputtering or
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`a directed vapor deposition process, the inner and outer
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`alumina layers having columnargrain structures.
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`25. A method according to claim 22, wherein the protec-
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`tive coating system is deposited on the thermal barrier
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`coating after the component has been removed from the gas
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`turbine engine and the thermal barrier coating has been
`cleaned.
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`26. A method according to claim 18, wherein the protec-
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`tive coating system is deposited on the thermal barrier
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`coating after polishing the thermal barrier coating-to-have a
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`surface finish of not greater than 0.75 micrometers Ra.
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`27. A method according to claim 18, wherein the inner
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`alumina layer is deposited to a thickness of about 0.5 to
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`about 50 micrometers, the platinum-group metal layer is
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`deposited to a thickness of about 0.1 to about 2 micrometers,
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`and the outer alumina layer is deposited to a thickness of
`about 0.5 to about 5 micrometers.
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`28. A method according to claim 18, further comprising
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`the step of depositing a layer of tantala on the outer alumina
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`layer.
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`29. A method according to claim 28, wherein the tantala
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`layer has a thickness of about 0.5 to about 2 micrometers.
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`30. A method of forming a protective coating system on
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`a thermal barrier coating of yttria-stabilized zirconia that is
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`present on a gas turbine engine component, the protective
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`10
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`15
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`20
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`35
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`40
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`45
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`US 6,627,323 B2
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`8
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`coating system defining an external surface of the
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`component, the method comprising the steps of:
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`depositing the inner alumina layer on the thermal barrier
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`coating so that the inner alumina layer has a columnar
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`grain structure;
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`depositing the platinum-group metal layer on the inner
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`alumina layer; and
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`depositing the outer aluminalayer on the platinum-group
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`layer so that
`the outer alumina layer has a
`metal
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