(12) United States Patent
`Nagaraj et al.
`
`USOO6627323B2
`(10) Patent No.:
`US 6,627,323 B2
`(45) Date of Patent:
`Sep. 30, 2003
`
`(54) THERMAL BARRIER COATING RESISTANT
`TO DEPOSITS AND COATING METHOD
`THEREFOR
`
`(56)
`
`(75) Inventors: Bangalore Aswatha Nagaraj, West
`Chester, OH (US); Jeffrey Lawrence
`Williams, Cincinnati, OH (US); John
`Frederick Ackerman, Laramie, WY
`(US)
`(73) Assignee: General Electric Company,
`Schenectady, NY (US)
`Subject to any disclaimer, the term of this
`patent is extended or adjusted under 35
`U.S.C. 154(b) by 12 days.
`
`(*) Notice:
`
`(21) Appl. No.: 10/079,036
`(22) Filed:
`Feb. 19, 2002
`(65)
`Prior Publication Data
`US 2003/0157361 A1 Aug. 21, 2003
`(51) Int. Cl. ............................ B32B 15/04; FO3B 3/12;
`C23C 16/00
`(52) U.S. Cl. ....................... 428/469; 428/697; 428/699;
`428/701; 428/702; 428/472; 428/336; 416/241 B;
`427/250; 427/255.19; 427/255.21; 427/255.23;
`427/255.31; 427/255.7; 204/192.16
`(58) Field of Search ................................. 428/469, 632,
`428/633, 701, 702, 697, 699, 472, 336;
`416/241 B; 427/255.19, 255.21, 255.23,
`255.31, 255.32, 255.34, 250, 255.7, 299,
`327; 204/192.16
`
`
`
`References Cited
`U.S. PATENT DOCUMENTS
`5,512.382 A * 4/1996 Strangman
`5,660,885 A 8/1997 Hasz et al.
`5,871,820 A 2/1999 Hasz et al.
`5,914,189 A 6/1999 Hasz et al.
`
`FOREIGN PATENT DOCUMENTS
`
`* 2/1999
`
`JP
`411O29380
`* cited by examiner
`Primary Examiner Deborah Jones
`ASSistant Examiner Jennifer McNeil
`(74) Attorney, Agent, or Firm-David L. Narciso; Gary M.
`Hartman; Domenica N. S. Hartman
`(57)
`ABSTRACT
`A protective coating System and method for protecting a
`thermal barrier coating from CMAS infiltration. The coating
`System comprises inner and Outer alumina layerS and a
`platinum-group metal layer therebetween. The outer alu
`mina layer is intended as a Sacrificial layer that reacts with
`molten CMAS, forming a compound with a melting tem
`perature Significantly higher than CMAS. AS a result, the
`reaction product of the 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 TBC, while the inner alumina
`layer appears to enhance the ability of the platinum-group
`metal layer to prevent CMAS infiltration.
`
`38 Claims, 1 Drawing Sheet
`
`Page 1 of 6
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`APPLIED MATERIALS EXHIBIT 1038
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`

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`U.S. Patent
`US. Patent
`
`Sep. 30, 2003
`Sep. 30, 2003
`
`US 6,627,323 B2
`US 6,627,323 B2
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`FG.
`FIG.
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`US 6,627,323 B2
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`1
`THERMAL BARRIER COATING RESISTANT
`TO DEPOSITS AND COATING METHOD
`THEREFOR
`
`CROSS REFERENCE TO RELATED
`APPLICATIONS
`
`Not applicable.
`
`STATEMENT REGARDING FEDERALLY
`SPONSORED RESEARCH
`Not applicable.
`
`BACKGROUND OF THE INVENTION
`
`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. Airplasma
`spraying (APS) has the advantages of relatively low equip
`ment costs and ease of application and masking, while
`TBC'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 coefficients of thermal
`expansion (CTE) 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 MCrAIX (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 (Al2O) 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 CTE 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
`grainstructures, respectively. Another detriment of CMAS is
`that the bond coat and substrate underlying the TBC 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, Silicides, 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.
`
`BRIEF SUMMARY OF THE INVENTION
`The present invention generally provides a protective
`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 CMAS, 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.
`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 MCrAIX 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
`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). ATBC 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.
`Of particular interest to the present invention is the
`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 (AlO4). 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. A suitable 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.
`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 CMAS 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 TBC 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 PVD 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
`protective coating System comprising inner and outer alu
`mina layers and a platinum-group metal layer encased
`therebetween.
`2. A component according to claim 1, wherein the thermal
`barrier coating is yttria-Stabilized Zirconia.
`3. A component according to claim 1, wherein the pro
`tective 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
`platinum-group metal layer consists essentially of platinum.
`5. A component according to claim 1, wherein the com
`ponent is an airfoil component of a gas turbine engine.
`6. A component according to claim 5, wherein the com
`ponent has a concave Surface, a conveX Surface and a leading
`edge therebetween, and the protective coating System over
`lies only one of the concave Surface, the conveX Surface or
`the 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 CMAS into the inner alumina layer.
`18. A method 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
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`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-group 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. A method according to claim 30, wherein CMAS has
`infiltrated the columnar grains of the outer alumina layer,
`and the platinum-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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`US 6,627,323 B2
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