`
`(cid:44)(cid:49)(cid:55)(cid:40)(cid:47) EXHIBIT 10(cid:22)(cid:27)
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`US. Patent
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`Sep. 30, 2003
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`US 6,627,323 B2
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`THERMAL BARRIER COATING RESISTANT
`TO DEPOSITS AND COATING METHOD
`THEREFOR
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`CROSS REFERENCE TO RELATED
`APPLICATIONS
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`Not applicable.
`STATEMENT REGARDING FEDERALLY
`SPONSORED RESEARCH
`
`Not applicable.
`BACKGROUND OF THE INVENTION
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`1. Field of the Invention
`
`This invention generally relates to coatings for compo-
`nents exposed to high temperatures, such as the hostile
`thermal environment of a gas turbine engine. More
`particularly, this invention is directed to a protective coating
`system for a thermal barrier coating on a gas turbine engine
`component, in which the protective coating system is resis-
`tant to infiltration by contaminants present in the operating
`environment of a gas turbine engine.
`2. Description of the Related Art
`Hot section components of gas turbine engines are often
`protected by a thermal barrier coating (TBC), which reduces
`the temperature of the underlying component substrate and
`thereby prolongs the service life of the component. Ceramic
`materials and particularly yttria-stabilized zirconia (YSZ)
`are widely used as TBC materials because of their high
`temperature capability, low thermal conductivity, and rela-
`tive ease of deposition by plasma spraying, flame spraying
`and physical vapor deposition (PVD) techniques. Air plasma
`spraying (APS) has the advantages of relatively low equip-
`ment costs and ease of application and masking, While
`'I‘IIC’s employed in the highest temperature regions of gas
`turbine engines are often deposited by PVD, particularly
`electron-beam PVD (EBPVD), which yields a strain-tolerant
`columnar grain structure. Similar columnar microstructures
`can be produced using other atomic and molecular vapor
`processes.
`To be effective, a TBC must strongly adhere to the
`component and remain adherent throughout many heating
`and cooling cycles. The latter requirement is particularly
`demanding due to the different coetficients of thermal
`expansion ((I'I‘L’) between ceramic materials and the sub-
`strates they protect, which are typically superalloys, though
`ceramic matrix composite (CMC) materials are also used.
`An oxidation-resistant bond coat is often employed to pro-
`mote adhesion and extend the service life of a TBC, as well
`as protect the underlying substrate from damage by oxida-
`tion and hot corrosion attack. Bond coats used on superalloy
`substrates are typically in the form of an overlay coating
`such as MCrAlX (where M is iron, cobalt and/or nickel, and
`X is yttrium or another rare earth element), or a diffusion
`aluminide coating. During the deposition of the ceramic
`TBC and subsequent exposures to high temperatures, such
`as during engine operation, these bond coats form a tightly
`adherent alumina (A1203) layer or scale that adheres the
`TBC to the bond coat.
`
`The service life of a TBC system is typically limited by
`a spallation event driven by bond coat oxidation and the
`resulting thermal fatigue. In addition to the (I'I‘L' mismatch
`between a ceramic TBC and a metallic substrate, spallation
`can be promoted as a result of the TBC being contaminated
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`with compounds found within a gas turbine engine during its
`operation. Notable contaminants include such oxides as
`calcia, magnesia, alumina and silica, which when present
`together at elevated temperatures form a compound referred
`to herein as CMAS. CMAS has a relatively low melting
`eutectic (about 1190° C.)
`that when molten is able to
`infiltrate to the cooler subsurface regions of a TBC, where it
`resolidifies. During thermal cycling,
`the CTE mismatch
`between CMAS and the TBC promotes spallation, particu-
`larly TBC deposited by PVD and APS due to the ability of
`the molten CMAS to penetrate their columnar and porous
`grain structures, respectively. Another detriment of CMAS is
`that the bond coat and substrate underlying the 'I‘IIC are
`susceptible to corrosion attack by alkali deposits associated
`with the infiltration of CMAS.
`
`Various studies have been performed to find coating
`materials that are resistant to infiltration by CMAS. Notable
`examples are U.S. Pat. Nos. 5,660,885, 5,871,820 and
`5,914,189 to Hasz et al., which disclose three types of
`coatings to protect a TBC from CMAS—related damage.
`These protective coatings are classified as being
`impermeable, sacrificial or non—wetting to CMAS. Imper—
`meable coatings are defined as inhibiting infiltration of
`molten CMAS, and include silica, tantala, scandia, alumina,
`hafnia, zirconia, calcium zirconate, spinels, carbides,
`nitrides, silicides, and noble metals such as platinum. Sac-
`rificial coatings are said to react with CMAS to increase the
`melting temperature or the viscosity of CMAS,
`thereby
`inhibiting infiltration. Suitable sacrificial coating materials
`include silica, scandia, alumina, calcium zirconate, spinels,
`magnesia, calcia and chromia. As its name implies, a non—
`wetting coating is non-wetting to molten CMAS, with
`suitable materials including silica, hafnia, zirconia, beryl—
`lium oxide, lanthana, carbides, nitrides, silicidcs, and noble
`metals such as platinum. According to the Hasz et al.
`patents, an impermeable coating or a sacrificial coating is
`deposited directly on the TBC, and may be followed by a
`layer of impermeable coating (if a sacrificial coating was
`deposited first), sacrificial coating (if the impermeable coat-
`ing was deposited first), or non—wetting coating. If used, the
`non-wetting coating is the outermost coating of the protec-
`tive coating system.
`While the coating systems disclosed by Hasz et al. are
`effective in protecting a TBC from damage resulting from
`CMAS infiltration, further improvements would be desir-
`able.
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`BRIEF SUMMARY OF THE INVENTION
`
`invention generally provides a protective
`The present
`coating system and method for protecting a thermal barrier
`coating (TBC) on a component used in a high—temperature
`environment, such as the hot section of a gas turbine engine.
`The invention is particularly directed to a protective coating
`system that significantly reduces if not prevents the infiltra-
`tion of CMAS into the underlying TBC.
`The protective coating system of this invention comprises
`inner and outer alumina layers and a platinum-group metal
`layer. The inner alumina layer is deposited on the thermal
`barrier coating, the platinum—group metal layer is deposited
`on the inner alumina layer, and the outer alumina layer is
`deposited on the platinum—group metal layer, so that the
`platinum-group metal layer is encased between the inner and
`outer alumina layers. The outer alumina layer is intended as
`a sacrificial layer that reacts with molten (XMAS, forming a
`compound with a melting temperature that is significantly
`higher than CMAS. As a result, the reaction product of the
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`outer alumina layer and CMAS resolidifies before it can
`infiltrate the TBC. The platinum-group metal
`layer
`is
`believed to serve as a barrier to infiltration of CMAS into the
`inner alumina layer and, therefore, the TBC. Notably, the
`inner alumina layer beneath the platinum-group metal layer
`appears to enhance the ability of the platinum-group metal
`layer to prevent infiltration of CMAS. In other words, the
`platinum—group metal layer is better able to perform as a
`barrier to CMAS infiltration if it is deposited on an alumina
`layer than if it were deposited directly on the TBC.
`In view of the above, the protective coating system of this
`invention is able to increase the temperature capability of a
`TBC by reducing the vulnerability of the TBC to spallation
`and the underlying substrate to corrosion from CMAS
`contamination. The layers of the protective coating system
`can be preferentially deposited on limited surface areas of a
`component more susceptible to CMAS contamination. In
`this manner, the additional weight and cost incurred by the
`protective coating system can be minimized. Finally,
`the
`protective coating system of this invention can be applied
`during the process of rejuvenating a TBC on a component
`returned from field service, thereby further extending the life
`of a TBC.
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`Other objects and advantages of this invention will be
`better appreciated from the following detailed description.
`BRIEF DESCRIPTION OF THE DRAWINGS
`
`FIG. 1 is a perspective View of a high pressure turbine
`blade.
`FIG. 2 is a cross-sectional view of the blade of FIG. 1
`along line 2—2, and shows a protective coating overlaying
`a thermal barrier coating in accordance with this invention.
`DETAILED DESCRIPTION OF THE
`INVENTION
`
`The present invention will be described in reference to a
`high pressure turbine blade 10 shown in FIG. 1, though the
`invention is generally applicable to any component
`that
`operates within a thermally and chemically hostile environ—
`ment. The blade 10 generally includes an airfoil 12 against
`which hot combustion gases are directed during operation of
`the gas turbine engine, and whose surfaces are therefore
`subjected to severe attack by oxidation, hot corrosion and
`erosion. The airfoil 12 is anchored to a turbine disk (not
`shown) with a dovetail 14 formed on a root section 16 of the
`blade 10. Cooling holes 18 are present in the airfoil 12
`through which bleed air is forced to transfer heat from the
`blade 10.
`
`The surface of the airfoil 12 is protected by a TBC system
`20, represented in FIG. 2 as including a metallic bond coat
`24 that overlies the surface of a substrate 22, the latter of
`which may be a superalloy and typically the base material of
`the blade 10. As Widely practiced with TBC systems for
`components of gas turbine engines, the bond coat 24 is
`preferably an aluminum-rich composition, such as an over-
`lay coating of an MCrAlX alloy or a diffusion coating such
`as a diffusion aluminide or a diffusion platinum aluminide,
`all of which are known in the art. Aluminum-rich bond coats
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`develop an aluminum oxide (alumina) scale 28, which is
`grown by oxidation of the bond coat 24. The alumina scale
`28 chemically bonds a TBC 26, formed of a thermal—
`insulating material, to the bond coat 24 and substrate 22. The
`TBC 26 of FIG. 2 is represented as having a strain-tolerant
`microstructure of columnar grains. As known in the art, such
`columnar microstructures can be achieved by depositing the
`TBC 26 using a physical vapor deposition (PVD) technique,
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`such as EBPVD. The invention is also applicable to nonco—
`lumnar TBC deposited by such methods as plasma spraying,
`including air plasma spraying (APS) A TBC of this type is
`in the form of molten “splats,” resulting in a microstructure
`characterized by irregular flattened grains and a degree of
`inhomogeneity and porosity.
`As with prior art TBC’s, the TBC 26 of this invention is
`intended to be deposited to a thickness that is sufficient to
`provide the required thermal protection for the underlying
`substrate 22 and blade 10. A suitable thickness is generally
`on the order of about 75 to about 300 micrometers. A
`preferred material for the TBC 26 is an yttria-stabilized
`zirconia (YSZ), a preferred composition being about 3 to
`about 8 weight percent yttria, though other ceramic mate-
`rials could be used, such as nonstabilized zirconia, or
`zirconia partially or fully stabilized by magnesia, ceria,
`scandia or other oxides.
`invention is the
`to the present
`Of particular interest
`susceptibility of TBC materials, including YSZ, to attack by
`CMAS. As discussed previously, CMAS is a relatively low
`melting eutectic that when molten is able to infiltrate colum-
`nar and porous TBC materials, and subsequently resolidify
`to promote spallation during thermal cycling. To address this
`concern, the TBC 26 in FIG. 2 is shown as being overcoated
`by a protective coating system 30 of this invention. As the
`outermost
`layer on the blade 10,
`the protective coating
`system 30 serves as a barrier to CMAS infiltration of the
`underlying TBC 26. The protective coating system 30 is
`shown in FIG. 2 as comprising four discrete layers 32, 34,
`36 and 38. The innermost layer 32 and the third layer 36 of
`the coating system 30 are formed of alumina (A1203). The
`layer 34 between the alumina layers 32 and 36 is formed of
`a platinum-group metal, which includes platinum,
`ruthenium, rhodium, palladium, osmium and iridium. The
`outermost layer 38 is an optional member of the coating
`system 30, and is intended to provide a nonstick surface to
`which CMAS will not readily wet and bond. A particularly
`suitable material for the outermost layer 38 is believed to be
`tantala, though it is foreseeable that other materials with
`similar nonstick properties could be used. A suitable thick-
`ness for the nonstick layer 38 is about 0.5 to about 5
`micrometers, more preferably about 0.5 to about 2 microme-
`ters.
`
`As represented in FIG. 2, the alumina layers 32 and 36
`have dense microstructures as a result of being deposited by
`PVD, chemical vapor deposition (CVD) or another suitable
`technique known in the art. The function of the inner and
`outer alumina layers 32 and 36 is to serve as sacrificial
`layers, reacting with molten CMAS that infiltrates the pro-
`tective coating system 30 to form one or more refractory
`phases with higher melting temperatures than CMAS. In
`effect, the alumina content of CMAS is increased above the
`eutectic point, yielding a modified CMAS with a higher
`melting and/or crystallization temperature. As a result, the
`reaction product of the inner and outer alumina layers 32 and
`36 and CMAS tends to resolidify before infiltrating the TBC
`26. Asuitable thickness for the outer alumina layer 36 is on
`the order of about 0.5 to about 5 micrometers, more pref-
`erably about 0.5 to about 2 micrometers, while a suitable
`thickness for the inner alumina layer 32 is believed to be
`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
`a barrier to infiltration of CMAS into the inner alumina layer
`32, thus enhancing the ability of the inner alumina layer 32
`to react with CMAS. A suitable method for depositing the
`metal layer 34 is again a CVD or PVD technique such as
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`sputtering. The platinum—group metal layer 34 is preferably
`entirely covered by the outer alumina layer 36, such that
`platinum-group metal is not present at the external surface of
`the coating system 30. With this arrangement,
`the outer
`alumina layer 36 serves to protect the platinum—group metal
`layer 34 from degradation. Importantly, the presence of the
`inner alumina layer 32 beneath the platinum—group metal
`layer 34 appears to enhance the ability of the platinum-group
`metal layer 34 to prevent infiltration of CMAS. In other
`words, improved resistant to (XMAS infiltration appears to
`be obtained if the platinum-group metal layer 34 is encased
`between the alumina layers 32 and 34, in comparison to a
`coating system in which the platinum-group metal layer is
`directly deposited on a TBC or is the outermost layer of the
`coating system. In its role as a barrier, a suitable thickness
`for the platinum—group metal layer 34 is believed to be about
`0.1 to about 2 micrometers, more preferably about 0.1 to
`about 0.5 micrometers. To promote the adhesion of the
`coating system 30, the surface of the TBC 26 is preferably
`polished prior to deposition of the inner alumina layer 32. A
`suitable surface finish is about 30 micro-inches (about 0.75
`micrometers) Ra or less.
`There are various opportunities for making use of the
`benefits of the protective coating system 30 of this invention.
`For example, the coating system 30 can be applied to newly
`manufactured components that have not been exposed to
`service. Alternatively, the coating system 30 can be applied
`to a component that has seen service, and whose TBC must
`be cleaned and rejuvenated before being returned to the
`field, In the latter case, applying the coating system 30 to the
`TBC can significantly extend the life of the component
`beyond that otherwise possible if the 'I‘IIC was not protected
`by the coating system 30. In a preferred embodiment, the
`coating system 30 is deposited only on those surfaces of a
`component that are particularly susceptible to damage from
`CMAS infiltration. In the case of the blade 10 shown in FIG.
`1, of particular interest
`is often the concave (pressure)
`surface 40 of the airfoil 12, which is can be significantly
`more susceptible to attack than the convex (suction) surface
`42 as a result of aerodynamic considerations. According to
`the invention, the layers 32, 34, 36 and optional layer 38 of
`the coating system 30 can be selectively deposited on the
`concave surface 40 of the airfoil 12, thus minimizing the
`additional weight and cost of the coating system 30. For this
`purpose, preferred deposition techniques include sputtering
`and directed PVD. Multiple blades can be simultaneously
`coated by positioning their convex surfaces back—to—back, so
`that their convex surfaces effectively mask each other and
`their concave surfaces face outward for coating. Deposition
`by sputtering or directed PVI) can then be performed to
`deposit the coating system 30 essentially exclusively on the
`exposed concave blade surfaces. While the concave surface
`40 of the airfoil 12 may be of particular interest, circum-
`stances may exist where other surface areas of the blade 10
`are of concern, such as the leading edge of the airfoil 12 or
`the region of the convex surface of the airfoil 12 near the
`leading edge.
`While the invention has been described in terms of a
`preferred embodiment, it is apparent that other forms could
`be adopted by one skilled in the art, such as by substituting
`other TBC, bond coat and substrate materials, or by utilizing
`other methods to deposit and process the protective coating
`system, Accordingly, the scope of the invention is to be
`limited only by the following claims.
`What is claimed is:
`1. A component having a thermal barrier coating on a
`surface thereof,
`the component comprising a protective
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`coating system overlying the thermal barrier coating, the
`wrotective coating system comprising inner and outer alu-
`mina layers and a platinum-group metal
`layer encased
`herebetween.
`2. Acomponent according to claim 1, wherein the thermal
`Jarricr coating is yttria-stabilized zirconia.
`3. A component according to claim 1, wherein the pro-
`ective coating system consists of the inner and outer alu—
`mina layers and the platinum-group metal layer.
`4. A component according to claim 1, wherein the
`filatinum-group metal layer consists essentially of platinum.
`5. A component according to claim 1, wherein the com-
`aonent is an airfoil component of a gas turbine engine.
`6. A component according to claim 5, wherein the com-
`aonent has a concave surface, a convex surface and a leading
`edge therebetween, and the protective coating system over-
`ies only one of the concave surface, the convex surface or
`he leading edge.
`7. A component according to claim 1, wherein the inner
`alumina layer has a thickness of about 0.5 to about 50
`micrometers, the platinum-group metal layer has a thickness
`of about 0.1 to about 2 micrometers, and the outer alumina
`layer has a thickness of about 0.5 to about 5 micrometers.
`8. A component according to claim 1, wherein the pro-
`tective coating system further comprises a layer of tantala
`overlying the outer alumina layer.
`9. A component according to claim 8, wherein the tantala
`layer has a thickness of about 0.5 to about 5 micrometers.
`10. A gas turbine engine component having a thermal
`barrier coating of yttria-stabilized zirconia, the component
`comprising an outer protective coating system overlying the
`thermal barrier coating, the protective coating system com-
`prising a platinum-group metal layer encased between inner
`and outer alumina layers having columnar grain structures,
`such that platinum—group metal is not present at an external
`surface of the component defined by the protective coating
`system.
`11. A component according to claim 10, wherein the
`protective coating system consists of the inner and outer
`alumina layers and the platinum-group metal layer, and the
`outer alumina layer defines the external surface of the
`component.
`12. A component according to claim 10, wherein the
`platinum—group metal layer consists essentially of platinum.
`13. A component according to claim 10, wherein the
`component
`is an airfoil component having a concave
`surface, a convex surface and a leading edge therebetween,
`and the protective coating system overlies only one of the
`concave surface, the convex surface or the leading edge.
`14. A component according to claim 10, wherein the inner
`alumina layer has a thickness of about 5 to about 10
`micrometers, the platinum-group metal layer has a thickness
`of about 0.1 to about 0.5 micrometers, and the outer alumina
`layer has a thickness of about 0.5 to about 2 micrometers.
`15. A component according to claim 10, wherein the
`protective coating system further comprises a layer of tan-
`tala overlying the outer alumina layer, and the tantala layer
`defines the external surface of the component.
`16. A component according to claim 15, wherein the
`tantala layer has a thickness of about 0.5 to about 2
`micrometers.
`17. A component according to claim 10, wherein CMAS
`has infiltrated the columnar grains of the outer alumina layer,
`the platinum-group metal layer being a barrier to infiltration
`of the (:MAS into the inner alumina layer.
`18. Amethod of protecting a thermal barrier coating on a
`surface of a component, the method comprising the step of
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`depositing a protective coating system on the thermal barrier
`coating, the protective coating system comprising an inner
`alumina layer deposited on the thermal barrier coating, a
`platinum-group metal layer deposited on the inner alumina
`layer, and an outer alumina layer deposited on the platinum-
`group metal layer so that the platinum-group metal layer is
`encased between the inner and outer alumina layers.
`19. A method according to claim 18, wherein the thermal
`barrier coating is yttria-stabilized zirconia.
`20. A method according to claim 18, wherein the protec—
`tive coating system consists of the inner and outer alumina
`layers and the platinum-group metal layer.
`21. A method according to claim 18, wherein the
`platinum-group metal layer consists essentially of platinum.
`22. A method according to claim 18, wherein the com-
`ponent is an airfoil component of a gas turbine engine.
`23. A method according to claim 22, wherein the com—
`ponent has a concave surface, a convex surface and a leading
`edge therebetween, and the protective coating system is
`selectively deposited on only one of the concave surface, the
`convex surface or the leading edge.
`24. A method according to claim 23, wherein each layer
`of the protective coating system is deposited by sputtering or
`a directed vapor deposition process, the inner and outer
`alumina layers having columnar grain structures.
`25. A method according to claim 22, wherein the protec—
`tive coating system is deposited on the thermal barrier
`coating after the component has been removed from the gas
`turbine engine and the thermal barrier coating has been
`cleaned.
`26. A method according to claim 18, wherein the protec-
`tive coating system is deposited on the thermal barrier
`coating after polishing the thermal barrier coating-to-have a
`surface finish of not greater than 0.75 micrometers Ra.
`27. A method according to claim 18, wherein the inner
`alumina layer is deposited to a thickness of about 0.5 to
`about 50 micrometers, the platinum—group metal layer is
`deposited to a thickness of about 0.1 to about 2 micrometers,
`and the outer alumina layer is deposited to a thickness of
`about 0.5 to about 5 micrometers.
`28. A method according to claim 18, further comprising
`the step of depositing a layer of tantala on the outer alumina
`layer.
`29. A method according to claim 28, wherein the tantala
`layer has a thickness of about 0.5 to about 2 micrometers.
`30. A method of forming a protective coating system on
`a thermal barrier coating of yttria-stabilized zirconia that is
`present on a gas turbine engine component, the protective
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`coating system defining an external surface of the
`component, the method comprising the steps of:
`
`depositing the inner alumina layer on the thermal barrier
`coating so that the inner alumina layer has a columnar
`grain structure;
`depositing the platinum—group metal layer on the inner
`alumina layer; and
`depositing the outer alumina layer on the platinum-group
`metal
`layer so that
`the outer alumina layer has a
`columnar grain structure,
`the platinum-group metal
`layer is encased between the inner and outer alumina
`layers, and platinum-group metal is not present at the
`external surface of the component.
`31. A method according to claim 30, wherein the protec-
`tive coating system consists of the inner and outer alumina
`layers and the platinum—goup metal layer, and the outer
`alumina layer defines the external surface of the component.
`32. A method according to claim 30, wherein the
`platinum-group metal layer consists essentially of platinum.
`33. A method according to claim 30, wherein the protec-
`tive coating system further comprises a layer of tantala
`deposited on the outer alumina layer so that the tantala layer
`defines the external surface of the component.
`34. Amethod according to claim 30, wherein CMAS has
`infiltrated the columnar grains of the outer alumina layer,
`and the platimun-group metal layer serves as a barrier to
`infiltration of the CMAS into the inner alumina layer.
`35. A method according to claim 30, wherein the com-
`ponent is an airfoil component having a concave surface, a
`convex surface and a leading edge therebetween, and the
`protective coating system is selectively deposited on only
`one of the concave surface, the convex surface or the leading
`edge.
`36. A method according to claim 35, wherein each layer
`of the protective coating system is deposited by sputtering or
`a directed vapor deposition process.
`37. A method according to claim 30, wherein the protec-
`tive coating system is deposited on the thermal barrier
`coating after the component has been removed from a gas
`turbine engine and the thermal barrier coating has been
`cleaned.
`38. A method according to claim 30, wherein the protec-
`tive coating system is deposited on the thermal barrier
`coating after polishing the thermal barrier coating to have a
`surface finish of not greater than 0.75 micrometers Ra.
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