Title:
PROTECTIVE COATING SYSTEMS FOR GAS TURBINE ENGINE APPLICATIONS AND METHODS FOR FABRICATING THE SAME
Kind Code:
A1


Abstract:
Protective coating systems for gas turbine engine applications and methods for fabricating such protective coating systems are provided. An exemplary method of manufacturing a turbine engine component includes providing a substrate in the form of the turbine engine component and forming a bond coating on and over the substrate. The method further includes forming a thermal barrier coating or an environmental barrier coating on and over the bond coating and forming a magnetoplumbite structure ceramic top coating on an over the thermal barrier coating or the environmental barrier coating.



Inventors:
Wali, Natalie (Chandler, AZ, US)
Application Number:
14/042323
Publication Date:
09/03/2015
Filing Date:
09/30/2013
Assignee:
HONEYWELL INTERNATIONAL INC.
Primary Class:
Other Classes:
60/772, 416/241R, 427/402, 427/419.1, 427/419.2, 427/446, 427/585, 427/596, 428/446, 428/448, 428/457, 428/697
International Classes:
C23C30/00; C23C28/04; F01D5/28; F01D25/00
View Patent Images:
Related US Applications:



Primary Examiner:
LAMBERT, WAYNE A
Attorney, Agent or Firm:
HONEYWELL/LKGLOBAL (Charlotte, NC, US)
Claims:
What is claimed is:

1. A method of manufacturing a turbine engine component comprising: providing a substrate in the form of the turbine engine component; forming a bond coating on and over the substrate; forming a thermal barrier coating or an environmental barrier coating on and over the bond coating; and forming a magnetoplumbite structure ceramic top coating on and over the thermal barrier coating or the environmental barrier coating.

2. The method of claim 1, wherein the turbine engine component is selected from the group consisting of: a turbine blade, a turbine shroud, a combustor, and a heat shield.

3. The method of claim 1, wherein providing the substrate comprises providing a substrate that is formed of a material selected from the group consisting of: a superalloy, a ceramic, and a ceramic matrix composite.

4. The method of claim 1, wherein forming the bond coating comprises forming a bond coating that comprises a material selected from the group consisting of: an aluminum-based material, a silicon-based material, a nickel-platinum aluminide material, an MCrAlY material, and an MCoCrAlY material.

5. The method of claim 1, wherein forming the thermal barrier coating or the environmental barrier coating comprises forming a thermal barrier coating or an environmental barrier coating that comprises a material selected from the group consisting of: yttria stabilized zirconia and yttria stabilized hafnia.

6. The method of claim 1, wherein forming the ceramic top coating comprises forming a ceramic top coating that comprises a material selected from the group consisting of: MAl12O19; MMeAl12O19; MAl11O19; MMeAl11O19; MAl11O18; and MMeAl11O18, where M and Me are cations selected independently from the group consisting of: Mg2+, La3+, Gd3+, Mn2+, Si4+, Na+, K+, Ca2+, Ba2+, La3+, Mg2+, Fe+, Fe3+, and Ti4+.

7. The method of claim 1, wherein forming the bond coating is performed using a method selected from the group consisting of: over-the-pack aluminizing, electroplating, electron beam physical vapor deposition, chemical vapor deposition, low pressure spraying, and cold spraying.

8. The method of claim 1, wherein forming the thermal barrier coating or the environmental barrier coating is performed using a method selected from the group consisting of: plasma spraying, physical vapor deposition, and electron beam physical vapor deposition.

9. The method of claim 1, wherein forming the ceramic top coating is performed using a method selected from the group consisting of: electron beam physical vapor deposition, air plasma spraying, high velocity oxy-fuel coating spraying, and sol-gel based methods.

10. The method of claim 1, further comprising installing the turbine engine component in a turbine engine.

11. A turbine engine component comprising: a substrate in the form of the turbine engine component; a bond coating on and over the substrate; a thermal barrier coating or an environmental barrier coating on and over the bond coating; and a magnetoplumbite structure ceramic top coating on and over the thermal barrier coating or the environmental barrier coating.

12. The turbine engine component of claim 11, wherein the turbine engine component is selected from the group consisting of: a turbine blade, a turbine shroud, a combustor, and a heat shield.

13. The turbine engine component of claim 11, wherein the substrate comprises a material selected from the group consisting of: a superalloy, a ceramic, and a ceramic matrix composite.

14. The turbine engine component of claim 11, wherein the bond coating comprises a material selected from the group consisting of: an aluminum-based material, a silicon-based material, a nickel-platinum aluminide material, an MCrAlY material, and an MCoCrAlY material.

15. The turbine engine component of claim 11, wherein the thermal barrier coating or the environmental barrier coating comprises a material selected from the group consisting of: yttria stabilized zirconia and yttria stabilized hafnia.

16. The turbine engine component of claim 11, wherein the ceramic top coating comprises a material selected from the group consisting of: MAl12O19; MMeAl12O19; MAl11O19; MMeAl11O19; MAl11O18; and MMeAl11O18, where M and Me are cations selected independently from the group consisting of: Mg2+, La3+, Gd3+, Mn2+, Si4+, Na+, K+, Ca2+, Ba2+, La3+, Mg2+, Fe+, Fe3+, and Ti4+.

17. A method of operating a turbine engine, the turbine engine comprising a turbine engine component comprising a substrate in the form of the turbine engine component, a bond coating on and over the substrate, a thermal barrier coating or an environmental barrier coating on and over the bond coating, and a magnetoplumbite structure ceramic top coating on an over the thermal barrier coating or the environmental barrier coating, the method comprising: exposing the turbine engine to a calcium-alumino-silicate (CMAS) material at a temperature sufficient such that at least a portion of the CMAS material is molten or becomes molten and adheres to the ceramic top coating of the turbine engine component.

18. The method of claim 17, wherein exposing the turbine engine comprises exposing the turbine engine to a temperature above about 2150° F.

19. The method of claim 18, further comprising cooling the turbine engine component below the temperature of about 2150° F. after the CMAS material has adhered to the ceramic top coating of the turbine engine component.

20. The method of claim 19, wherein exposing the turbine engine comprises exposing the turbine engine to combustion gasses during the operation of the gas turbine engine.

Description:

TECHNICAL FIELD

The present disclosure generally relates to coatings for gas turbine engine applications and methods for fabricating such coatings.

BACKGROUND

Turbine engines are used as the primary power source for various kinds of aircraft and other vehicles. The engines may also serve as auxiliary power sources that drive air compressors, hydraulic pumps, and industrial electrical power generators. Most turbine engines generally follow the same basic power generation procedure. Compressed air is mixed with fuel and burned, and the expanding hot combustion gases are directed against stationary turbine vanes in the engine. The vanes turn the high velocity gas flow partially sideways to impinge onto turbine blades mounted on a rotatable turbine disk. The force of the impinging gas causes the turbine disk to spin at high speed. Jet propulsion engines use the power created by the rotating turbine disk to draw more air into the engine, and the high velocity combustion gas is passed out of the gas turbine aft end to create forward thrust. Other engines use this power to turn one or more propellers, electrical generators, or other devices.

Because fuel efficiency increases as engine operating temperatures increase, turbine engine blades and vanes are typically fabricated from high-temperature materials such as nickel-based superalloys. However, although nickel-based superalloys have good high-temperature properties and many other advantages, they may be susceptible to corrosion, oxidation, thermal fatigue, and/or foreign particle impact when exposed to harsh working environments during turbine engine operation. As a result, diverse coatings and coating methods have been developed to increase the operating temperature limits and service lives of the high pressure turbine components, including the turbine blade and vane airfoils.

The durability and the maximum temperature capability of a coating system used in gas turbine engines are often limited by deposits of partially or fully molten sand, often referred to as calcia-magnesia-alumina-silicate (CMAS) for the major oxide constituents. CMAS deposits may degrade ceramic coatings by several mechanisms. In one example related superalloy components that are coated with ceramic thermal barrier coatings, CMAS may deposit on top of the ceramic coatings. The first CMAS phase to melt, known as the incipient melt, may then infiltrate the coating porosity and be drawn through the coating thickness via capillary forces. Upon engine cooldown this infiltrant will freeze and form a high modulus phase. Since the thermal barrier coatings rely on spatially configured voids to achieve strain tolerance with the superalloy substrate, those regions penetrated by the CMAS can be detrimental, causing the thermal barrier coating to be susceptible to extensive spallation when subjected to subsequent thermal cycles. This loss of strain compliance may lead to early spallation. In another degradation mechanism, the CMAS deposit may react with the ceramic coating material. The solid crystalline reaction product might also lead to loss of strain compliance. Thermal barrier coating spallation can lead to a drastic reduction in the turbine engine component durability and to a direct attack on the underlying (superalloy) substrate.

Accordingly, it is desirable to provide protective coating systems for gas turbine engine applications that exhibit long life and high reliability. It also is desirable to provide protective coating systems that minimize or eliminate spalling due to CMAS deposits. It is further desirable to provide methods for fabricating such protective coating systems. Furthermore, other desirable features and characteristics of the present invention will become apparent from the subsequent detailed description of the invention and the appended claims, taken in conjunction with the accompanying drawings and this background of the invention.

BRIEF SUMMARY

The present disclosure generally relates to coatings for gas turbine engine applications and methods for fabricating such coatings. In one embodiment, an exemplary method of manufacturing a turbine engine component includes providing a substrate in the form of the turbine engine component and forming a bond coating on and over the substrate. The method further includes forming a thermal barrier coating or an environmental barrier coating on and over the bond coating and forming a magnetoplumbite structure ceramic top coating on an over the thermal barrier coating or the environmental barrier coating.

In another embodiment, an exemplary turbine engine component includes a substrate in the form of the turbine engine component, a bond coating on and over the substrate, a thermal barrier coating or an environmental barrier coating on and over the bond coating, and a magnetoplumbite structure ceramic top coating on an over the thermal barrier coating or the environmental barrier coating.

In yet another embodiment, an exemplary method of operating a turbine engine is provided. The turbine engine used in this method includes a component including a substrate in the form of the turbine engine component, a bond coating on and over the substrate, a thermal barrier coating or an environmental barrier coating on and over the bond coating, and a magnetoplumbite structure ceramic top coating on an over the thermal barrier coating or the environmental barrier coating. The method includes exposing the turbine engine to a calcium-alumino-silicate (CMAS) deposit at a temperature sufficient such that at least a portion of the CMAS material is molten or becomes molten and adheres to the ceramic top coating or infiltrates into the ceramic top coating of the turbine engine component.

This summary is provided to introduce a selection of concepts in a simplified form that are further described below in the detailed description. This summary is not intended to identify key features or essential features of the claimed subject matter, nor is it intended to be used as an aid in determining the scope of the claimed subject matter.

BRIEF DESCRIPTION OF THE DRAWINGS

The present invention will hereinafter be described in conjunction with the following drawing figures, wherein like numerals denote like elements, and wherein:

FIG. 1 is a schematic illustration of a gas turbine blade upon which an exemplary embodiment of a protective coating system of the present disclosure may be disposed;

FIG. 2 is a cross-sectional view of a protective coating system in accordance with an exemplary embodiment of the present disclosure;

FIGS. 3-5 are cross-sectional views of a protective coating system, such as the protective coating system of FIG. 2, illustrating its resistance of the coating system to spallation by deposited CMAS;

FIG. 6 is a flowchart of a method for fabricating a protective coating system, such as the protective coating system of FIG. 2, in accordance with an exemplary embodiment of the present disclosure; and

FIG. 7 is a flowchart of a method of operation of a turbine engine component using a protective coating system, such as the protective coating system of FIG. 2, in accordance with an exemplary embodiment of the present disclosure.

DETAILED DESCRIPTION

The following detailed description of the invention is merely exemplary in nature and is not intended to limit the invention or the application and uses of the invention. Furthermore, there is no intention to be bound by any theory presented in the preceding background of the invention or the following detailed description of the invention.

The present disclosure includes a protective coating system for a variety of substrates, including gas turbine and aero-engine components. As is generally appreciated in the art, gas turbine engines that operate in environments with airborne particles such as sand and volcanic ash are prone to the accumulation of siliceous deposits in the hot sections. When these adherent particles become molten or partially molten they are referred to as “CMAS” for the predominant phases of calcia (CaO), alumina (Al2O3), magnesia (MgO), and silica (SiO2). CMAS is benign for most uncoated metallic-based components, however, it is generally a life-limiting factor for thermal barrier coatings (TBC) and environmental barrier coatings (EBC). TBC's are typically applied to the metallic surfaces of combustors, heat shields, turbine shrouds, and turbine blades. During operation, TBC surfaces are often exposed to hot gases at temperatures above about 2200° F., for example. At these high temperatures, portions of CMAS melt/wick into the pores or vertical cracks (for plasma-sprayed) or intercolumnar gaps (for electron beam vapor phase deposited) of the TBC (or EBC). The depth of penetration is dependent on many factors, including time, melt viscosity, and the temperature gradient in the TBC (or EBC). Reactions between the melt and the coating may also occur.

Upon cool-down during engine shutoff (i.e., to a temperature below the above-noted melting temperature), the CMAS freezes in the open porosity of the TBC/EBC. This is particularly detrimental since these cracks and pores provide the TBC/EBC with strain compliance to relieve the stresses that results from the coefficient of thermal mismatch between the coating and the substrate. The CMAS inhibits this property so that the TBC/EBC cracks to relieve the stresses that build up over multiple cycles. This process continues through engine startup and shutdown cycles until the TBC/EBC spalls. Creating a CMAS resistant coating is complicated by the fact that sand has no universal composition. Though calcia, magnesia, alumina, and silica are major components, there may also be oxides of titanium, iron, potassium, and sodium.

The present disclosure provides a ceramic material with a layered crystal structure that may serve as a CMAS mitigation strategy to protect the underlying TBC and/or EBC. This ceramic topcoat may serve to not only arrest CMAS infiltration/penetration of the TBC and/or EBC, but may also act as a sacrificial topcoat that gradually cleaves away with continued CMAS deposition and/or reaction. For use as this sacrificial topcoat, β-alumina magnetoplumbite structured materials possess a particularly attractive crystal structure that lends itself to cleavage much like mica. Magentoplumbites are composed of layered spinel blocks nominally of the composition [X11O16+], where X is a cation, stabilized by cations and oxygen anions in mirror planes between spinel layers. To illustrate the latter, some magnetoplumbite structures may be represented as M+O[X11O16] or M3+O2[X11O16] where the portion in brackets is the spinel block and the portion outside the brackets represents the interspinel constituents. An attractive aspect of this structure is that although the spinel block remains constant, the number of atoms that may be located in the interspinel layer may vary. The number and combination of atoms that can fit in the interspinel layer may be adjusted until the close-packed structure of M2+XO3[X11O16] is achieved. In this case, the M2+ ion is coordinated and in a nearly close-packed oxygen lattice. Mechanisms of charge balance are possible in either the interspinel layer, between the interspinel layer and spinel block, or in the spinel block itself For example, in magnetoplumbite compounds of the form M2+X12O19, the M2+ ion can be substituted with a trivalent ion of about the same size, or there may be mixed occupancy of the site. Though the magnetoplumbite structure is stable for a large number of compositions and stoichiometries, for the purposes of this disclosure only those materials that are stable at elevated temperatures are considered.

For example, magnetoplumbites and beta-alumina-type oxides are such hexaaluminates that are stable at elevated temperatures, and as such are suitable for use as the topcoat of the present disclosure. They possess a variety of attractive characteristics that make them suitable as a CMAS-mitigation topcoat, including, for example: 1) a high aspect ratio crystalline phase that resists high temperature densification; 2) a crystal structure composed of easily cleavable crystallographic planes; 3) the ability to accommodate a wide variety of cations without disrupting the crystal structure; and 4) large fracture energy anisotropy between the planes parallel and normal to the basal planes that can impart toughness that other state-of-the-art CMAS resistant coatings lack in practical use (e.g. gadolinium zirconate and delta phase zirconia). Embodiments of the present disclosure are set forth in greater detail below with reference to the Figures.

FIG. 1 illustrates a superalloy blade 150 that is exemplary of the types of components or substrates that are used in turbine engines, although turbine blades commonly have different shapes, dimensions and sizes depending on gas turbine engine models and applications. However, this invention is not restricted to such substrates and may be utilized on many other substrates requiring thermal barrier protection, including other components of gas turbine engines exposed to high temperature gases. Nickel-based superalloys are one class of materials that are for use in manufacturing turbine engine blades, and other classes of materials including cobalt-based superalloys, titanium-based superalloys, nickel aluminides including NiAl, alumina fiber/alumina silicate matrix composites, silicon carbide fiber/silicon carbide matrix composites, alumina fiber/refractory metal matrix composites, alumina fiber/MCrAlY matrix composites, refractory metal fiber/MCrAlY matrix composites, alumina fiber/NiAl matrix composites, silicon carbide fiber/gamma TiAl matrix composites, refractory metal fiber/NiAl matrix composites, carbon fiber/carbon matrix composites, alumina fiber/TiAl alloy matrix composites, silicon carbide fiber/alumna matrix composites, silicon carbide fiber/silicon nitride matrix composites and other materials systems are suitable as well. In a particular embodiment, the substrate blade is formed of a superalloy, a ceramic, or a ceramic matrix composite material. The illustrated blade 150 has an airfoil portion 152 including a pressure surface 153, an attachment or root portion 154, a leading edge 158 including a blade tip 155, and a platform 156. The blade 150 may be formed with a non-illustrated outer shroud attached to the tip 155. The blade 150 may have non-illustrated internal air-cooling passages that remove heat from the turbine airfoil. After the internal air has absorbed heat from the superalloy, the air is discharged into a combustion gas flow path through passages 159 in the airfoil wall.

FIG. 2 is a cross-sectional view of a substrate 10 upon which is disposed a protective coating system 12 in accordance with an exemplary embodiment of the present disclosure. The substrate 10 may be, for example, a turbine blade such as turbine blade 150 of FIG. 1. The protective coating system 12 overlies the substrate 10 and any intermediate layers, and is formed of a bond coating 14, a thermal and/or environmental barrier coating 18, and a top coat 22 formed of a layered crystallographic structure magnetoplumbite ceramic, as initially noted above. In one exemplary embodiment, the bond coating 14 is a simple diffusion aluminide. In another embodiment, the bond coating 14 is a more complex diffusion aluminide that includes another layer such as another metal layer. In one embodiment, the other metal layer is a platinum layer. In another exemplary embodiment, the bond coating 14 is an overlay coating known as an MCrAlX coating, wherein M is cobalt, iron, and/or nickel, or an oxidation resistant intermetallic, such as diffusion aluminide, platinum aluminide, or an active element-modified aluminide. In some bond coatings, the chromium may be omitted. The X is hafnium, zirconium, yttrium, tantalum, rhenium, ruthenium, palladium, platinum, silicon, titanium, boron, carbon, or combinations thereof. Some examples of MCrAlX compositions include NiAlCrZr and NiAlZr. In a particular embodiment, the bond coating 14 is Al-based, either a nickel-platinum aluminide (which can be modified with Hf additions) or an MCrAlY/MCoCrAlY bond coat, where M is a metal, such as Fe. For a ceramic or ceramic matrix composite, the bond coat can be any Si-based material, including pure silicon, silicate or silicide.

Thermal and/or environmental barrier coating 18 may include, for example, a stabilized zirconia-based thermal barrier coating, such as yttria stabilized zirconia (YSZ), or a stabilized hafnia-based thermal barrier coating, such as yttria stabilized hafnia (YSH). In a particular embodiment, the thermal barrier coating or environmental barrier coating may be a multilayer system, whether functionally graded (e.g., porosity level) or composition (e.g., 23YSZ/8YSZ).

The topcoat 22 includes a layered crystallographic structure oxide ceramic. The final topcoat composition is nominally any of the following compositions: MAl12O19; MMeAl12O19; MAl11O19; MMeAl11O19; MAl11O18; MMeAl11O18; where M and Me are cations including the cations listed here: Mg2+, La3+, Gd3+, Mn2+, Si4+, Na+, Ca2+, Ba2+, La3+, Mg2+, Fe+, Fe3+, or Ti4+. This list of cations is exemplary and those having ordinary skill in the art will readily envision other cations that may be used as well, based on this disclosure. Cations as small as Na+ and K+ can be immobilized by the hexaaluminates. Since these salts have been found to penetrate down to the bond coat, presumably via water vapor, and so contribute to coating spallation, the arrest of their penetration is another attractive feature of a magnetoplumbite coating. As such, the topcoat is expected to provide additional life to the underlying TBC and EBC by preventing CMAS-induced spallation. Additional information regarding the preparation and characterization of magnetoplumbite materials may be found in the following seven references, each of which is incorporated by reference herein in its entirety: 1. P E D Morgan and E H Cirlin, Comm Am Ceram Soc C-114-C-115 (1982); 2. M K Cinibulk Ceram Eng &Sci Proc 721-728 (1994); 3. C Friedrich, R Gadow and T Schimer J Therm Spray Technol 10(4) 592-598 (2001); 4. L Haoran, W Chang-An, Z Chenguang J Am Ceram Soc 96 [4] 1063-1066 (2013); 5. B Jiang, M H Fang, Z H Huang, Y G Liu, P Peng, and J Zhang Mater Res Bull 45 1506-1508 (2010); 6. N P Bansal, D Zhu and M Eslamloo-Grams NASA/TM-2007-214850 (2007); 7. R Collongues, D Gourier, A Kahn-Harari, A M Lejus, J Thery, and D Vivien Annu Rev Mater Sci 20 51-82 (1990).

Having described the general structure of the protective coating system 12, a method 30 for fabricating a protective coating system, such as protective coating system 12 of FIG. 2, will now be described. Referring to FIG. 6, the method 30 begins with the step of providing a substrate 10 (step 32). As described above, the substrate may be a turbine blade, or any other turbine component such as, for example, a vane or a shroud, that is subjected to high gas temperatures. The substrate may include nickel-based superalloys, cobalt-based superalloys, titanium-based superalloys, nickel aluminides, including NiAl, and any of the other materials or material systems discussed above for the fabrication of substrate 10 of FIG. 2. A bond coating, such as bond coating 14 of FIG. 2, then is formed on the substrate (step 34). The bond coating may include any of the materials described above for bond coating 14. The bond coating may be deposited using various known deposition techniques such as, for example, simple over-the-pack aluminizing, electroplating, electron beam physical vapor deposition (EB-PVD), chemical vapor deposition (CVD), low pressure spray, and cold spraying and may be deposited to a thickness, indicated by double-headed arrow 15, in the range of about 25 μm (about 1 mil) to about 2000 μm (about 8 mils). After formation, the exposed surface of the bond coating may be cleaned, such as by grit blasting, to remove any oxides or contaminants that have formed on or adhered to the bond coating surface.

The method continues with the formation of a thermal barrier coating or environmental barrier coating, such as thermal barrier coating or environmental barrier coating 18 of FIG. 2 (step 36). In one exemplary embodiment, the thermal barrier coating is yttria stabilized zirconium oxide (YSZ) that is deposited by plasma spraying, PVD or EB-PVD. In another exemplary embodiment, the thermal barrier coating is yttria stabilized hafnium oxide (YSH) that is deposited by plasma spraying or EB-PVD. A thickness of thermal barrier coating 18 may vary according to design parameters and may be, for example, between about 50 and about 1000 μm, and typically between about 100 and about 300 μm.

The method continues with the formation of the topcoat included of the layered crystallographic structure oxide ceramic (step 38). It may be applied by EB-PVD, plasma spray or via liquid precursor routes, such as sol-gel or coating with a starting slurry. The topcoat material may be prepared by blending commercially available ceramic powders in appropriate proportions, such as Al2O3, M oxides, and Me oxides. The powders mixed in proper ratio and may be ball-mill mixed with arabic gum and deionized water and spay-dried to prepare free-flowing powders that are suitable for use in the above-noted applications methods.

Reference is now made to FIGS. 3-5, which illustrate a cross-sectional view of a turbine engine component in accordance with FIG. 2 during operation, and FIG. 7, which is a flowchart setting forth a method 40 of operating a turbine engine in accordance with the illustrations in FIGS. 3-5. With particular reference to FIG. 3, during operation of the gas turbine engine, sand or other CMAS materials 50 will deposit and adhere to the topcoat 22 as partially or fully molten deposits (step 41). At a critical elevated temperature, such as above about 2150° F., there will be an incipient melting phase 24 (i.e., the first phase of CMAS to melt) (step 43). Upon melting, the cations in the CMAS melt 24 will be captured by the magnetoplumbite topcoat (now referred to as reference numeral 22a to indicate a change in composition based on CMAS infiltration) and stabilized in the crystal structure, thus arresting the infiltration of the melt by “freezing” the remainder of the deposit (step 45). With reference now to FIG. 4, as the solid CMAS deposit 26 grows in thickness, stresses will develop in the top or topmost layers of the hexaaluminate coating 22a (step 47). Finally, with reference to FIG. 5, at a critical stress these CMAS-covered layers will cleave off and expose underlying fresh topcoat layers (now referred to as reference numeral 22b to indicate fresh topcoat layers exposed by the cleaving process) (step 49). This process will repeat and continue until the topcoat is removed, at which point the underlying TBC or EBC 18 is exposed.

Accordingly, protective coating systems for gas turbine engine applications and methods for fabricating such protective coating systems have been provided. The systems include a magnetoplumbite structure ceramic topcoat that provides additional life to the underlying TBC and EBC by preventing CMAS-induced spallation.

While at least one exemplary embodiment has been presented in the foregoing detailed description of the invention, it should be appreciated that a vast number of variations exist. It should also be appreciated that the exemplary embodiment or exemplary embodiments are only examples, and are not intended to limit the scope, applicability, or configuration of the invention in any way. Rather, the foregoing detailed description will provide those skilled in the art with a convenient road map for implementing an exemplary embodiment of the invention, it being understood that various changes may be made in the function and arrangement of elements described in an exemplary embodiment without departing from the scope of the invention as set forth in the appended claims and their legal equivalents.