HOT CORROSION-RESISTANT COATINGS FOR GAS TURBINE COMPONENTS

Abstract

A gas turbine component for use in a gas turbine engine includes a substrate a ceramic-based thermal barrier coating (TBC), and a diffusion chromide bond coat between the base material and the TBC. A thermally grown oxide (TGO) layer can be formed on the bond coat prior to application of the TBC. The TBC and the TGO include a common metal oxide. The oxide can be sacrificially in use and soluble in a molten sulfate salt, make the coating system particularly suitable for use in a marine environment.

Description

TECHNICAL FIELD

This disclosure generally relates to coatings and surface treatments for gas turbine components.

BACKGROUND

Certain critical gas turbine components, particularly those in the hot section of a gas turbine, are exposed to a harsh operating environment that may expose the component to high temperatures, high mechanical stresses, and potentially reactive combustion gases. The possible effects of this type of operating environment may be considered when selecting turbine component materials. For example, material characteristics such as resistance to heat, stress, fatigue, corrosion, erosion, and/or oxidation may be considered. Material costs and manufacturability may be considered as well, along with numerous other factors. While many advancements have been made in turbine component materials and coatings in the aviation industry, they have failed to address certain aspects of the operating environment of non-aviation applications, such as marine vessel powerplant applications. In such applications, airborne salts contribute to corrosion, and peak operating temperatures may be sustained for longer periods than in applications in which the turbine is primarily used for propulsion. Exposure of the turbine component to turbine fuels, some of which include contaminants such as vanadium or sulfur, can also contribute to corrosion.

These types of conditions will likely cause turbine components in future naval gas turbine engines to have reduced service life if not addressed, for example from about 20,000 hours currently to about 10,000 hours or less, due to elevated temperatures and the effects of sodium and calcium sulfate deposition on turbine component surfaces. This could lead to shorter intervals for engine removal and have a significant impact on the readiness of the U.S. Navy surface fleet while also significantly increasing costs (e.g., maintenance, repair, replacement, etc.). These next generation gas turbines may use thermal barrier coatings (TBCs) on turbine blades and vanes to improve engine fuel efficiency and increase horsepower. In addition, the engines may need to operate at higher power levels to support electrical generation for other shipboard uses. The effect will be an increased percentage of time at peak temperatures compared to current levels and/or increased turbine component temperatures even with the use of TBCs to reduce component substrate temperatures.

SUMMARY

In accordance with various embodiments, a gas turbine component for use in a gas turbine engine includes a substrate comprising a metal base material, a ceramic-based thermal barrier coating disposed over the substrate, and a bond coat disposed between the base material and the thermal barrier coating. The thermal barrier coating defines at least a portion of an outer surface of the gas turbine component, and the bond coat includes a chromide diffusion coating.

In some embodiments, the bond coat comprises a platinum diffusion coating.

In some embodiments, the bond coat comprises an aluminide diffusion coating.

In some embodiments, the bond coat further comprises hafnium, silicon, zirconium, or any combination thereof.

In some embodiments, the bond coat is a diffusion coating comprising chromium, platinum, aluminum, and at least one of hafnium, silicon or zirconium.

In some embodiments, the base material is a Ni-based superalloy having a gamma phase and a gamma prime phase distributed within the gamma phase. Platinum of the diffusion coating resides in the gamma prime phase, and chromium of the diffusion coating resides in the gamma phase.

In some embodiments, the component is a gas turbine blade comprising an airfoil that is exposed to combustion gases of a gas turbine engine when in use. The bond coat and the thermal barrier coating are located along the airfoil.

In accordance with various embodiments, a method of making a gas turbine component having a thermal barrier coating defining at least a portion of an outer surface of the gas turbine component includes the step of forming a diffusion bond coat on a component substrate. The diffusion bond coat includes chromium interdiffused with a metal substrate material of the component substrate, and the thermal barrier coating is subsequently coated over the bond coat.

In some embodiments, the diffusion bond coat comprises a Pt-aluminide coating.

In some embodiments, the step of forming the diffusion bond coat comprises vapor phase deposition of the chromium on the metal substrate material.

In some embodiments, the diffusion bond coat comprises hafnium, silicon, zirconium, or any combination thereof.

In some embodiments, the step of forming the diffusion bond coat comprises the steps of coating a slurry comprising a platinum-group metal over the substrate material and heat treating the slurry-coated substrate material to interdiffuse the platinum-group metal with the substrate material.

In some embodiments, the slurry further comprises hafnium, silicon, zirconium, or any combination thereof

In some embodiments, the step of forming the diffusion bond coat further comprises vapor phase deposition of the chromium on the metal substrate material before the step of coating the slurry over the substrate material.

In some embodiments, the slurry comprises the chromium of the bond coat, and the chromium is interdiffused with the substrate material during the step of heat treating.

In some embodiments, the method includes vapor phase aluminide coating the component substrate after the step of heat treating, whereby the diffusion bond coat further comprises an aluminide coating.

In some embodiments, the method includes forming a thermally grown oxide layer over the diffusion bond coat, and the thermal barrier coating is subsequently coated over the thermally grown oxide layer.

In accordance with various embodiments, a gas turbine component for use in a gas turbine engine includes a substrate comprising a metal base material, a metal bond coat formed on the base material, a thermally grown oxide layer formed on the bond coat and comprising an oxide of a metal element of the bond coat, and a ceramic-based thermal barrier coating disposed over the substrate and defining at least a portion of an outer surface of the gas turbine component. The thermal barrier coating further comprises the oxide of the metal element of the bond coat.

In some embodiments, the oxide of the metal element is alumina or chromia.

In some embodiments, the bond coat includes a diffusion coating comprising chromide, Pt-aluminide, and at least one of hafnium, silicon, or zirconium.

In accordance with various embodiments, a method of making a gas turbine component includes the step of coating a turbine component substrate with a coating system that includes a ceramic-based thermal barrier coating having a sacrificial oxide that is soluble in a molten sulfate salt.

It is contemplated that the various features set forth in the preceding paragraphs, in the claims and/or in the following description and drawings may be taken independently or in any combination thereof. For example, features disclosed in connection with one embodiment are applicable to all embodiments, except where there is incompatibility of features.

BRIEF DESCRIPTION OF THE DRAWINGS

Illustrative embodiments will hereinafter be described in conjunction with the appended drawings, wherein like designations denote like elements, and wherein:

FIG. 1 is a perspective view of an exemplary gas turbine component that may include the coating system described herein;

FIG. 2 is a cross-sectional view of a portion of a gas turbine component illustrating an embodiment of the coating system comprising a diffusion bond coat;

FIG. 3 is a cross-sectional view of a portion of a gas turbine component illustrating an embodiment of the coating system comprising an overlay bond coat;

FIG. 4 is a process flow diagram illustrating an exemplary method of coating a gas turbine component substrate; and

FIG. 5 is a process flow diagram illustrating another exemplary method of coating a gas turbine component substrate.

DETAILED DESCRIPTION

The coating system and methods described herein address multiple problems that are likely to be encountered with turbine components as operating temperatures increase, particularly in a marine environment and particularly in applications in which the turbine operates at or near peak power levels a majority of the time.

In the marine environment, airborne sea salt can be ingested by the turbine engine and, even though filtration systems may reduce the level of sea salt, sufficient levels may be present to condense sulfate salts onto some surfaces. The dew point for condensation of sodium sulfate is dependent on pressure and salt concentration but is generally below about 1500-1700° F. Conventional turbine materials with conventional hot corrosion resistant coatings can survive for about 20,000 hours under such conditions, even without a thermal barrier coating (TBC). For next-generation gas turbines, one might expect that higher operating temperatures would cause turbine components to remain above the dew point of sodium sulfate such that less sodium sulfate would condense and less sulfidation-type hot corrosion would occur, leaving oxidation resistance as the main challenge.

However, it has been found that deposition of such salts will occur during low power operation, when even TBC surfaces are below the dew point of sulfate salts. When present on the surface of a TBC, the molten salt can wick into the porosity of the coating or into the columnar structure of certain TBCs (e.g., those applied via electron beam physical vapor deposition) and penetrate to the thermally grown oxide (TGO) layer on the underlying bond coat. Once penetrated into the vacancies of the TBC, these salts may have a very long residence time even when temperatures later rise above their dew point. Thus, the rate of sulfidation-type corrosion beneath the TBC can increase, and the TBC life to spallation may drop to unacceptably low levels. When a portion of the TBC is lost to spallation, component temperatures increase locally at the spalling site, and hot corrosion/oxidation rates will be excessive, leading to early engine removals.

Another sulfate salt that may be present in the marine environment is calcium sulfate, the presence of which may modify the corrosion environment and indirectly contribute to more severe (e.g., acidic) conditions. This condition has been observed in under-platform locations on turbine blades and may well become a factor in future airfoil TBC applications. The coating system described herein addresses this problem, among others.

Another failure mechanism in turbine components involves impact damage at the leading edge of turbine blade airfoils. TBC's are somewhat susceptible to local damage from large particle (50 μm or larger) impact. This can lead to local loss of the TBC accompanied by a local rise in underlying component temperature. In aircraft engines, the time at maximum operating temperature is relatively low because only a small percentage of overall operating time (e.g., during aircraft acceleration for take-off) is at peak power levels. In such cases, conventional bond coats can often protect the component substrate at the local spall site for its full service live. This will not be the case in aero-derivative gas turbines that operate at full power most of the time. In these applications, local spalls in the TBC can lead to burn through, even with some level of de-rated power. Next-generation marine gas turbines would thus benefit from improved TBC impact resistance and/or a more graceful failure mode, as made possible by the coating system described herein.

Further, there is a trend toward more reliance on the coating system to protect the underlying substrate material, with next-generation marine gas turbines likely using 2^nd generation single crystal alloys that have much lower hot corrosion resistance than conventional superalloys (e.g., Rene 80). Should the bond coat be penetrated, whether by impact damage or spallation from corrosion beneath the TBC or some other cause, such substrates may rapidly succumb to hot corrosion attack.

The coating system described below may be used on gas turbine blades or other gas turbine components, such as compressor blades, turbine or compressor vanes, seals, rotors or hubs, shafts, shrouds, blisks, or any other component that may encounter the harsh environment present in a gas turbine or other type of combustion engine. The coating system includes a thermal barrier coating (TBC) and an underlying bond coat that together function to improve hot corrosion resistance of the coated component. The coating system is particularly suitable for use in marine or other applications where airborne salts are sometimes ingested by the turbine where they react with other combustion constituents and/or fuel contaminants and/or become molten. While described in the context of a marine gas turbine engine intended to power waterborne naval vessels, for example, the coating system may be employed in other applications while realizing the same or other advantages.

Referring to FIG. 1, an exemplary gas turbine component is shown. In this embodiment, the component is a gas turbine blade 10. The turbine blade 10 includes an airfoil 12 and a shank 14, each extending from opposite sides of a platform 16. The airfoil 12 may include a cross-section or profile configured to cause high and low pressure regions on opposite sides thereof when placed in a particular orientation in a flowing fluid, thus causing the blade 10 to move in the direction of lower pressure. The shank 14 may be used as part of an attachment to secure the blade to a hub or other component that rotates about a central axis. The shank 14 may include several features not individually described here, such as a root, neck, ridges, sealing flanges, “angel wings,” etc. In operation, multiple blades 10 may be arranged so that the airfoils extend radially away from the central axis of the hub to form a turbine that can transform energy from axial gas flow into rotational motion, or vice versa where the blade is used as part of a compressor or turbine.

The platform 16 lies between the airfoil 12 and the shank 14, generally dividing the blade into an upper boundary region 18 and a lower boundary region 20. The upper boundary region 18, also referred to as the gas path region, is exposed to combustion gases during operation and includes the airfoil 12 and a topside 22 of the platform 16. The lower boundary region 20 is generally not exposed to combustion gases during operation and includes an underside 24 of the platform 16 and any other blade components under the platform or on the shank side of the platform. This arrangement of turbine blade components may also result in the lower boundary region 20 operating at temperatures lower than those at which the gas path region 18 operates. For example, the lower boundary region 20 may operate at temperatures that range from about 1200-1600° F., while the gas path region 18 may operate at higher temperatures that may range from about 1900-2100° F. For purposes of this disclosure, the gas path region 18 may also be referred to as the airfoil portion, and the lower boundary region 20 may be referred to as the shank portion.

In the illustrated embodiment, the turbine blade 10 also includes internal cooling channels 26, the ends of some of which are shown along the airfoil surface. Channels 26 may extend from one or more surfaces of the shank portion 20 to one or more surfaces of the airfoil portion 18 to facilitate the flow of a cooling fluid such as air therethrough. Various blade cooling arrangements are known in the art, and the cooling channels may be omitted entirely in some cases.

Due to the harsh environment in and around an operating gas turbine engine, engine components are sometimes constructed using superalloy materials that have high strength, ductility, and creep resistance at high temperatures and relatively high resistance to corrosion. Superalloy materials may be based on nickel (Ni), cobalt (Co), and/or Iron (Fe). Examples of superalloys include alloys available under the trade names Hastelloy, Inconel, and René, such as René N4, René N5, or others. While the corrosion-resistance of superalloys may generally be considered very good as metal alloys are concerned, the elevated temperatures and stresses, corrosive combustion gases, and other elements (e.g., atmospheric pollutants or particulates, fuel additives and impurities, salts, etc.) in the gas turbine operating environment can accelerate the corrosion of even the most corrosion-resistant superalloys. As used herein, corrosion may be divided into two categories, including Type I and Type II corrosion. Type I corrosion generally refers to relatively high temperature corrosion that occurs above about 1600° F. and includes sulfidation-type corrosion. Type II corrosion generally refers to relatively low temperature corrosion that occurs below about 1600° F.

Various types of coatings or surface treatments have been developed in attempt to improve the corrosion and fatigue resistance of superalloy components in gas turbines. One approach includes the use of a thermal barrier coating (TBC), particularly along the airfoil portion of turbine blades. TBCs are generally designed to thermally insulate the underlying materials from the high temperature combustion gases so that less thermal energy is absorbed by the turbine blade, thereby allowing the superalloy or other substrate material to operate at a lower temperature than would otherwise be possible. In some cases, the gas turbine component can operate in an environment in which the temperature of the combustion gases exceeds the melting point of the substrate material when a suitable TBC is employed.

Referring to FIGS. 2 and 3, partial cross-sections of exemplary gas turbine components 10 are shown, including a gas turbine component substrate 30, such as a blade substrate, and a coating system 40 that includes a thermal barrier coating 42. The substrate 30 may be formed from a base material 32 by casting and/or other known processing techniques. The base material 32 may be a metal or metal-based alloy capable of forming a gamma/gamma prime (γ/γ′) microstructure in which the gamma prime phase is in the form of a precipitate distributed within the gamma phase matrix. One example of such an alloy is a Ni-based superalloy that can form a gamma/gamma prime microstructure when heat-treated under certain conditions. For example, heat-treating a Ni-based superalloy can cause a gamma prime phase to form as a precipitate that includes Ni₃Al and/or Ni₃Ti distributed in a gamma phase that is a solid solution including Ni and other elements. Any of the above-mentioned exemplary superalloys, as well as other alloys capable of forming a gamma/gamma prime microstructure, may be suitable for use as the base material 32. Cobalt-based superalloys may also be suitable. The substrate 30 may also include materials or layers of materials other than the base material 32. For example, substrate 30 may include a layer of the base material 32 clad or otherwise attached to a different underlying material.

The coating system 40 includes the thermal barrier coating 42 along with one or more other coatings or material layers that overlie the base material 32 and provides an increased resistance to corrosion relative to the base material, particularly at high temperatures and/or in an operating environment that includes airborne salts, such as sulfate salts. The coating system 40 of FIG. 2 includes a diffusion bond coat 44 at an outer surface 46 of the substrate 30 and an oxide layer 48 between the bond coat 44 and the TBC 42. The coating system 40 of FIG. 3 includes the same layers as the system of FIG. 2, except the bond coat 44′ is an overlay coating rather than a diffusion coating.

Each coating layer may be classified as either an overlay coating or a diffusion coating. Both types of coatings may be at least partially interdiffused with an underlying material, but any interdiffusion that is present with an overlay coating is in the form of a relatively thin layer at the interface of the overlay coating and the underlying material. A diffusion coating has a substantial portion of its thickness interdiffused with the underlying material, and may be entirely interdiffused with the underlying material. Diffusion coatings may be designated with an “-ide” suffix and indicate the element or elements interdiffused with the underlying material. For example, aluminide and chromide coatings are diffusion coatings of aluminum and chromium, respectively. Thus, a diffusion coating is a layer of material that is richer in the coated and diffused element(s) than the underlying material and further includes the constituent elements of the underlying coated material.

The thermal barrier coating 42 is an overlay coating and, as noted above, is generally configured to thermally insulate the substrate 30 from the high combustion gas temperatures in operation. Suitable thermal barrier coatings may be ceramic-based materials, such as yttria-stabilized zirconia (YSZ) or a rare earth zirconate. The TBC may be coated over the substrate 30 and bond coat by known methods.

The TBC 42 may be porous, have a specific microstructure (e.g., columnar voids), or have cracks or microcracks oriented perpendicular to the outer surface of the TBC and thereby allow corrosive elements such as oxygen or other substances to migrate through the TBC toward the substrate 30. One of the functions of the bond coat 44 is to provide corrosion protection for the substrate material 32. Corrosion of the substrate material 32 beneath the TBC 42 can be doubly problematic because it not only affects the integrity of the substrate material, but it can also cause spalling of the TBC due to the stresses induced in the TBC by the growth of the underlying corrosion. When TBC spalling occurs, the thermal barrier properties of the TBC are lost locally and the underlying material is exposed to the high temperature combustion gases, accelerating any corrosion and reducing fatigue life.

The oxide layer 48 is not necessarily present during turbine blade fabrication. The oxide layer 48 may be a thermally grown oxide (TGO) that forms above about 1300° F. from oxidation of the material (e.g., the bond coat 44) beneath the TBC 42, through which oxygen and other gases passes through during operation due to TBC porosity. In some embodiments of the coating system 40 described herein, the oxide layer 48 is pre-grown or otherwise formed prior to application of the TBC 42. As discussed below, controlling the composition of the oxide layer 48 has been found to be advantageous, particularly with respect to sulfidation-type corrosion. In addition, the TGO has been found to greatly enhance adhesion of the TBC to the bond coat 44.

In the embodiment of FIG. 2, the bond coat 44 may include a chromide diffusion coating. While chromide diffusion coatings have been proposed for use in turbine components to improve corrosion resistance, applications have been limited to lower temperature portions of such components, such as the shank portion of a turbine blade. Chromide diffusion coatings have generally been considered insufficient on higher temperature regions of turbine components, such as the airfoil portion of a turbine blade. However, inclusion of chromide in the bond coat 44 has now been found to offer certain advantages. In particular, the presence of chromide in the bond coat 44 ensures the presence of chromia in the oxide layer 48, which behaves differently from the alumina that is present in the oxide layer when aluminide is present in the bond coat, particularly with respect to sulfidation-type corrosion.

For example, it has been determined that a molten film of sodium sulfate on a substrate surface without a TBC will flux (i.e., dissolve) a protective alumina film formed thereon. However, alumina does not saturate the molten salt in a manner that prevents further fluxing. Instead, the alumina (or aluminate) reprecipitates near the outer surface of the film of molten salt, and fluxing continues. Chromia behaves differently. While chromia may also dissolve in the molten salt, it reprecipitates at the lower oxygen-activity location within the molten salt—i.e., near the protective oxide layer. In this manner, the presence of chromia in the oxide layer 48 of the illustrative coating system 40 offers better Type I corrosion protection and, in particular, better resistance to sulfidation-type corrosion. The presence of the chromide diffusion coating in the bond coat 44 can help ensure the presence of chromia in the TGO 48 and has thus unexpectedly been found to offer improved high temperature corrosion resistance when used as part of the bond coat 44 beneath the TBC 42.

The chromide diffusion coating can be formed in a number of ways, each including deposition of chromium on the substrate surface to be coated and interdiffusion of the chromium with the substrate material by heat treatment. In one embodiment, the chromide diffusion coating is formed by vapor phase deposition, in which the substrate is exposed to a vapor containing the chromium to be deposited. Vapor phase deposition is typically performed in a chamber environment at a temperature sufficient to diffuse the chrome into the substrate surface immediately upon deposition such that separate deposition and heat treatment steps are not required. For example, the same elevated temperature required for diffusion can be used to heat a bed containing chromium metal and an activator such that chromium-containing vapor is produced, with the vapor coming into contact with the substrate and depositing the chromium on the substrate, which may be supported over the bed vapor source. The chromide diffusion coating may also be formed in a slurry process, in which the chromium is deposited on the substrate surface as part of a coating comprising the chromium, an activator, and a liquid carrier that allows the coating to be coated onto the substrate by spraying, painting, dipping, etc., with the liquid component being evaporated away prior to diffusion by heat treating. The chromide diffusion coating may also be formed in a pack cementation process, in which the chromium is deposited on the substrate surface by packing the substrate surface in a bed (e.g., a powder bed) of material containing chromium and activator, similar to the above-described vapor phase deposition bed. In both the pack cementation and slurry processes, chromium-containing vapor may be produced from the respective bed of material and slurry coating during the heat treatment step. But these processes differ from vapor phase deposition in that at least some of the chromium metal is in contact with the substrate before heat treating begins and during heat treating. Also, other constituents can be added to the slurry coating mixture or to the pack-cementation bed to be diffused into the substrate. Other deposition methods can be used as well, such as electroplating, PVD, or other suitable techniques, followed by diffusion by heat treatment.

In one embodiment, the chromide diffusion coating is formed by combining a powder contact process and a simultaneous vapor phase process, resulting in a super-chromide diffusion coating, at least a portion of which has a chromium content of 60 wt % or higher. In one example, the portion of the substrate to be coated is packed into a bed of powder material comprising chromium metal and an activator. Then, while heating the substrate in the bed of material, a chromium-containing vapor is flowed through the bed of material from a vapor source outside of the bed of material to boost the concentration of chrome-containing vapor that the packed portion of the substrate is exposed to. The result is a chromide diffusion coating with a higher concentration of chromium than can be produced by any other chromide coating technique alone. Super-chromide coatings and methods of producing them are discussed further in U.S. Patent Application Publication No. 2013/0115907, hereby incorporated by reference.

Where the bond coat 44 is a diffusion coating as illustrated in FIG. 2, it may also include one or more of the following elements interdiffused with the substrate material: a platinum-group metal such as platinum (Pt), aluminum (Al), hafnium (Hf), silicon (Si), or zirconium (Zr).

Platinum and aluminum can be combined in diffusion coatings in the form of a bond coat as a Pt-aluminide coating. Conventionally, a Pt-aluminide bond coat is formed on a substrate by first electroplating the Pt onto the substrate surface, then using vapor phase deposition to deposit the aluminum on the part surface while the high temperature environment of the vapor phase process diffuses both the Pt and the aluminum into the substrate surface to form the Pt-aluminide diffusion coating. Alternatively, the electroplated Pt may be heat treated in a vacuum to diffuse it into the substrate prior to the vapor phase aluminide process.

Where the substrate material is a metal or metal-based alloy capable of forming a gamma/gamma prime (γ/γ′) microstructure as described above, the Pt or other platinum-group metal can be interdiffused with the substrate material in a separate step from the aluminide coating to form a gamma/gamma prime diffusion coating. In this type of diffusion coating, the Pt-group metal resides in the gamma prime phase of the heat treated component. Such coatings have recently found uses as corrosion-resistant coatings separately from their utility as bond coats for TBCs. For example, U.S. Pat. No. 9,297,089, incorporated herein by reference, describes such coatings used on the lower temperature shank portion of turbine blades in the absence of aluminide and TBCs. This type of coating has also been found to perform well as a bond coat for a TBC, but it can be susceptible to hot corrosion and oxidation failure quickly if the TBC coating is compromised.

In one embodiment, the bond coat 44 comprises a platinum-group diffusion coating. In a particular embodiment, the platinum-group diffusion coating is formed by a slurry process similar to that described above in conjunction with the diffusion chromide coating. For example, a platinum-group metal such as Pt may be deposited on the substrate surface as part of a slurry coating comprising the Pt-group metal, an activator, and a liquid carrier that allows the coating to be coated onto the substrate by spraying, painting, dipping, etc., with the liquid component being evaporated away prior to diffusion by heat treating. Application by the slurry process offers certain advantages over electroplating, which has long been the preferred method of applying Pt in a Pt-aluminide coating. For example, other elements can be added to the slurry mixture to be simultaneously applied and subsequently diffused into the substrate material together with the Pt-group metal. In a particular embodiment, the above-described chromide diffusion coating is formed in the same slurry process as the Pt-group diffusion coating by including both chromium and the Pt-group metal in the slurry mixture. The slurry process also facilitates selective deposition of the coating constituents. For example, the slurry can be sprayed or otherwise coated onto the turbine component substrate only where desired, while electroplating typically results in the deposition of platinum over the entire conductive substrate surface, thereby wasting an expensive coating constituent.

In other embodiments, a dopant is added to the Pt-group slurry, such as hafnium (Hf), silicon (Si), or zirconium (Zr). These dopants are diffused into the substrate material as part of the bond coat 44 and act to stabilize the TGO layer 48 to enhance TBC adherence.

The bond coat 44 may further include an aluminide diffusion coating, which may be formed via vapor phase deposition and simultaneous heat treatment to diffusion the aluminum into the substrate material. When the aluminide diffusion coating is formed after the Pt-group metal is already interdiffused with the substrate material, the result is a complex Pt-aluminide coating that is somewhat different from a Pt-aluminide coating that is formed when electroplated Pt and aluminum are simultaneously diffused into the substrate material. For example, the pre-diffused Pt-group metal already resides in a gamma prime phase of the coating before the aluminum is diffused into the already existing coating. The aluminide forms a particularly complex Pt-aluminide when the Pt-group diffusion coating is formed via a slurry process in which the slurry includes a TGO stabilizing dopant (e.g., hafnium).

FIG. 4 illustrates an exemplary method 100 of forming the coating system according to a specific embodiment. The illustrated method includes the step 110 of providing a turbine component substrate, the step 120 of forming a diffusion bond coat over at least a portion of the substrate, and the step 130 of forming a thermal barrier coating that provides the outer surface of the finished turbine component. In this example, the diffusion bond coat is formed via steps 140, 150, and 160, including respectively forming a chromide diffusion coating, a doped Pt-group slurry diffusion coating, and an aluminide diffusion coating on the substrate. Step 150 is performed via a slurry process as described above in which the slurry mixture includes the Pt-group metal (e.g., platinum) along with one or more TGO-stabilizing dopants, such as Hf, Si, or Zr. Each of steps 140 and 160 may be performed via a vapor phase deposition process as described above or by other suitable processes such as a pack cementation process or separate slurry processes. In one specific embodiment, step 150 includes step 140 with the chromium being included as part of the doped-Pt slurry mixture.

In another specific embodiment, step 140 is performed in a combined process in which the chromium is provided by two different sources, including a powder source in contact with the substrate (e.g., pack cementation) and a vapor source separate from the powder source (e.g., chrome-containing vapor flowed through the pack cementation bed), resulting in a super-chrome diffusion coating before the slurry coat and aluminide coating. The doped Pt-group diffusion coating may then be formed by applying a dopant-containing Pt-slurry over the super-chrome coated substrate and diffusing the Pt-group metal and dopants into the chrome-coated surface. This results in a gamma/gamma prime system in which the Pt-group metal resides in the gamma prime phase of the heat treated substrate material with the chrome residing in the gamma phase (e.g., with the nickel component of the substrate material in solid solution). A vapor phase aluminide diffusion coating is then formed over the gamma/gamma prime system, resulting in a complex Pt-aluminide coating. After completion of the step 120 of forming the bond coat, the TBC is applied thereover in step 130. The TBC may be a ceramic or ceramic-based coating such as yttria-stabilized zirconia (YSZ) or other suitable material and may be applied via an air plasma process, solution plasma spray (SPS), electron-beam physical vapor deposition (EB-PVD), or other suitable process.

The complex Pt-aluminide coating thus formed results in the dopants being located in specific locations of the bond coat where it can be beneficial to the TGO and/or TBC layers. More particularly, the dopants (e.g., Hf, Si, and/or Zr) will be distributed throughout the thickness of the aluminide portion of the bond coat, with some of the dopant near the aluminide-coated surface and some of the dopant diffused farther into the substrate. In contrast, adding such dopants during the vapor phase aluminide process after more conventional Pt electroplating may result in aggregation of the dopants at the outermost portion of the aluminide as a localized hafnide that is detrimental to formation of the TGO layer where present at the bond coat surface, which negatively affects TBC adhesion. In order to be beneficial, dopants such as hafnium must be present in lower concentrations and more evenly distributed within the thickness of the bond coat. It is desirable to have the dopants a few microns (e.g., 1-3 μm) from the TGO and below the surface of the bond coat. This is very difficult to achieve with Pt electroplate followed by a doped aluminide vapor phase process.

The example of FIG. 4 also illustrates an optional step 170 in which a thermally grown oxide layer is formed on the bond coat prior to application of the thermal barrier coating. In another example, the TGO layer can be formed after application of the TBC but before the finished turbine component is put into service. In either case, pre-forming the TGO layer offers the advantage of forming the oxide under controlled conditions rather than allowing it to form under the unpredictable conditions present during initial operation of the component.

FIG. 5 illustrates another exemplary method 200 of forming the coating system according to a specific embodiment. The illustrated method includes the step 210 of providing a turbine component substrate, the step 220 of forming a bond coat over at least a portion of the substrate, and the step 230 of forming a doped thermal barrier coating that provides the outer surface of the finished turbine component. In this example, the diffusion bond coat is formed via step 250 of forming an overlay bond coat (e.g., MXCrAly) by known methods. Optionally, a chrome diffusion coating is formed in step 240 prior to application of the overlay bond coat.

In this example, the TBC is doped to include a sacrificial oxide which will dissolve in and at least partially saturate any molten salts, such as sodium sulfate, that makes its way into the porosity of the TBC. Suitable dopants include alumina or chromia which will dissolve in the molten salt and thus slow or halt fluxing of the TGO layer. The dopant may be applied with the TBC in a thermal spray process in an amount of about 1 wt % to 3 wt %. The example of FIG. 5 also illustrates an optional step 270 in which the TGO layer is formed on the bond coat prior to application of the thermal barrier coating. In another example, the TGO layer can be formed after application of the TBC but before the finished turbine component is put into service.

It is to be understood that the foregoing is a description of one or more preferred exemplary embodiments of the invention. The invention is not limited to the particular embodiment(s) disclosed herein, but rather is defined solely by the claims below. Furthermore, the statements contained in the foregoing description relate to particular embodiments and are not to be construed as limitations on the scope of the invention or on the definition of terms used in the claims, except where a term or phrase is expressly defined above. Various other embodiments and various changes and modifications to the disclosed embodiment(s) will become apparent to those skilled in the art. All such other embodiments, changes, and modifications are intended to come within the scope of the appended claims.

As used in this specification and claims, the terms “for example,” “e.g.,” “for instance,” “such as,” and “like,” and the verbs “comprising,” “having,” “including,” and their other verb forms, when used in conjunction with a listing of one or more components or other items, are each to be construed as open-ended, meaning that that the listing is not to be considered as excluding other, additional components or items. Other terms are to be construed using their broadest reasonable meaning unless they are used in a context that requires a different interpretation.

Claims

1. A gas turbine component for use in a gas turbine engine, comprising:

a substrate comprising a metal base material;

a ceramic-based thermal barrier coating disposed over the substrate, the thermal barrier coating defining at least a portion of an outer surface of the gas turbine component; and

a bond coat disposed between the base material and the thermal barrier coating, the bond coat comprising a chromide diffusion coating.

2. A gas turbine component as defined in claim 1, wherein the bond coat comprises a platinum diffusion coating.

3. A gas turbine component as defined in claim 1, wherein the bond coat comprises an aluminide diffusion coating.

4. A gas turbine component as defined in claim 1, wherein the bond coat further comprises hafnium, silicon, zirconium, or any combination thereof.

5. A gas turbine component as defined in claim 1, wherein the bond coat is a diffusion coating comprising chromium, platinum, aluminum, and at least one of hafnium, silicon or zirconium.

6. A gas turbine component as defined in claim 5, wherein the base material is a Ni-based superalloy having a gamma phase and a gamma prime phase distributed within the gamma phase, the platinum of the diffusion coating residing in the gamma prime phase and the chromium of the diffusion coating residing in the gamma phase.

7. A gas turbine component as defined in claim 1, wherein the component is a gas turbine blade comprising an airfoil that is exposed to combustion gases of a gas turbine engine when in use, the bond coat and the thermal barrier coating being located along the airfoil.

8. A method of making a gas turbine component having a thermal barrier coating defining at least a portion of an outer surface of the gas turbine component, the method comprising the step of forming a diffusion bond coat on a component substrate, the diffusion bond coat comprising chromium interdiffused with a metal substrate material of the component substrate, wherein the thermal barrier coating is subsequently coated over the bond coat.

9. The method of claim 8, wherein the diffusion bond coat comprises a Pt-aluminide coating.

10. The method of claim 8, wherein the step of forming the diffusion bond coat comprises vapor phase deposition of the chromium on the metal substrate material.

11. The method of claim 8, wherein the diffusion bond coat comprises hafnium, silicon, zirconium, or any combination thereof.

12. The method of claim 8, wherein the step of forming the diffusion bond coat comprises the steps of coating a slurry comprising a platinum-group metal over the substrate material and heat treating the slurry-coated substrate material to interdiffuse the platinum-group metal with the substrate material.

13. The method of claim 12, wherein the slurry further comprises hafnium, silicon, zirconium, or any combination thereof.

14. The method of claim 12, wherein the step of forming the diffusion bond coat further comprises vapor phase deposition of the chromium on the metal substrate material before the step of coating the slurry over the substrate material.

15. The method of claim 12, wherein the slurry comprises the chromium of the bond coat, the chromium being interdiffused with the substrate material during the step of heat treating.

16. The method of claim 12, further comprising the step of vapor phase aluminide coating the component substrate after the step of heat treating, whereby the diffusion bond coat further comprises an aluminide coating.

17. The method of claim 8, further comprising the step of forming a thermally grown oxide layer over the diffusion bond coat, wherein the thermal barrier coating is subsequently coated over the thermally grown oxide layer.

18. A gas turbine component for use in a gas turbine engine, comprising:

a substrate comprising a metal base material;

a metal bond coat formed on the base material;

a thermally grown oxide layer formed on the bond coat and comprising an oxide of a metal element of the bond coat; and

a ceramic-based thermal barrier coating disposed over the substrate and defining at least a portion of an outer surface of the gas turbine component, wherein the thermal barrier coating further comprises the oxide of the metal element of the bond coat.

19. A gas turbine component as defined in claim 18, wherein the oxide of the metal element is alumina or chromia.

20. A gas turbine component as defined in claim 18, wherein the bond coat comprises a diffusion coating comprising chromide, Pt-aluminide, and at least one of hafnium, silicon, or zirconium.

21. A method of making a gas turbine component comprising the step of coating a turbine component substrate with a coating system comprising a ceramic-based thermal barrier coating comprising a sacrificial oxide that is soluble in a molten sulfate salt.