System and Method for Rejuvenating Coated Components of Gas Turbine Engines

The present disclosure is directed to a method for rejuvenating a damaged coated component of a gas turbine engine. The method includes uninstalling the damaged coated component from the gas turbine engine. The method also includes isolating a first coated portion of the component of the gas turbine engine from a second coated portion of the component. In addition, the method includes simultaneously depositing a first coating material on the first coated portion of the component and a different, second coating material on the second coated portion of the component. The method also includes reinstalling the rejuvenated coated component into the gas turbine engine.

FIELD OF THE INVENTION

The present invention relates generally to gas turbine engines, and more specifically, to systems and methods for rejuvenating coated components, such as turbine blades, by simultaneously depositing multiple coatings thereon.

BACKGROUND OF THE INVENTION

A gas turbine engine generally includes, in serial flow order, a compressor section, a combustion section, a turbine section and an exhaust section. In operation, air enters an inlet of the compressor section where one or more axial or centrifugal compressors progressively compress the air until it reaches the combustion section. Fuel is mixed with the compressed air and burned within the combustion section to provide combustion gases. The combustion gases are routed from the combustion section through a hot gas path defined within the turbine section and then exhausted from the turbine section via the exhaust section.

In particular configurations, the turbine section includes, in serial flow order, a high pressure (HP) turbine and a low pressure (LP) turbine. The HP turbine and the LP turbine each include various rotatable turbine components such as a rotor shaft, rotor disks mounted or otherwise carried by the rotor shaft, turbine blades mounted to and radially extending from the periphery of the disks, and various stationary turbine components such as stator vanes or nozzles, turbine shrouds, and engine frames. The rotatable and stationary turbine components at least partially define the hot gas path through the turbine section. For example, the gas turbine buckets or blades generally have an airfoil shape designed to convert the thermal and kinetic energy of the flow path gases into mechanical rotation of the rotor. As the combustion gases flow through the hot gas path, thermal energy is transferred from the combustion gases to the rotatable and stationary turbine components.

Turbine blades may be constructed of a number of superalloys (e.g., nickel-based superalloys), as well as ceramic matrix composites coated with an environmental barrier coating to avoid oxidation and recession in the presence of high temperature steam during operation of the engine. Current coating processes for engine components include a two-step process that first includes coating a first portion of the component with a corrosion-resistant coating (such as a chromide coating) and then subsequently includes coating a second portion of the component with an oxidation-prohibiting coating, such as an aluminide coating. For example, for high-pressure turbine blades, the shank is first coated via a pack process using a powder chromide coating and a later step includes coating the airfoil of the turbine blade with an aluminide coating.

One of the issues associated with conventional two-step coating techniques, however, includes chloride gases leaking from the pack during the chromide coating process which can damage the airfoil. More specifically, in engine components with internal aluminide coatings, the depletion of aluminum due to reaction with chloride gases can result in low engine performance and/or a high scrap rate.

In addition, during operation of the gas turbine engine, the coatings originally applied to the turbine blades begin to wear off due to oxidation and other environmental conditions within the turbine engine. Conventional repair methods for turbine blades require “full repair” of the affected blades, which includes uninstalling the turbine blade from the engine, stripping the previously applied coatings therefrom, and repeating the two-step process described above. Thus, full repair of the turbine blades can be time-consuming and expensive.

In view of the aforementioned, an improved system and method for rejuvenating coated turbine blades rather than requiring full repair of such blades would be advantageous. More specifically, a system and method for rejuvenating turbine blades by simultaneously depositing multiple coatings on the previously-coated blade would be desired in the art.

BRIEF DESCRIPTION OF THE INVENTION

In one aspect, the present disclosure is directed to a method for rejuvenating a damaged coated component of a gas turbine engine. The method includes uninstalling the damaged coated component from the gas turbine engine. The method also includes isolating a first coated portion of the component of the gas turbine engine from a second coated portion of the component. In addition, the method includes simultaneously depositing a first coating material on the first coated portion of the component and a different, second coating material on the second coated portion of the component. The method also includes reinstalling the rejuvenated coated component into the gas turbine engine.

In another aspect, the present disclosure is directed to a kit for rejuvenating a damaged coated component of a gas turbine engine. The kit includes a first coating material and a second coating material, wherein the first and second coating materials are different. The kit also includes a maskant for isolating a first coated portion of the component of the gas turbine engine from a second coated portion of the component and a coating system configured to simultaneously deposit the first and second coating materials on the first and second coated portions of the component, respectively.

In yet another aspect, the present disclosure is directed to a method for rejuvenating a damaged coated turbine blade of a gas turbine engine. The method includes uninstalling the damaged coated turbine blade from the gas turbine engine. The method also includes placing a coated shank of the turbine blade in a masking chamber having a masking lid so as to isolate the shank from a coated airfoil of the turbine blade. Another step includes filling a masking chamber with a chromium-based powder coating. Thus, the method also includes simultaneously depositing an aluminum-based coating on the airfoil via diffusion coating.

DETAILED DESCRIPTION OF THE INVENTION

Chemical elements are discussed in the present disclosure using their common chemical abbreviation, such as commonly found on a periodic table of elements. For example, aluminum is represented by its common chemical abbreviation Al; chromium is represented by its common chemical abbreviation Cr; and so forth.

Generally, the present disclosure is directed to an improved system and method for rejuvenating damaged coated components of gas turbine engines by simultaneously depositing multiple coatings at different locations on the component. More specifically, in certain embodiments, the method includes uninstalling the damaged coated component from the gas turbine engine and isolating a first coated portion of the component from a second coated portion of the component. For example, the component may include a turbine blade having a coated airfoil and a shank that may or may not be coated. Thus, the method includes simultaneously depositing a first coating material on the first coated or uncoated portion (i.e. the shank) of the component and a different, second coating material on the second coated portion (i.e. the airfoil) of the component. The method also includes reinstalling the rejuvenated coated component into the gas turbine engine.

As such, even though the chromium-based powder and the aluminum-based material are deposited onto the turbine blade simultaneously, the coating materials do not contact each other or neighboring portions of the blade. Thus, the present disclosure provides a robust process that reduces time and costs associated with coating engine components. Further, by depositing the coating materials at the same time, degradation of the coating materials is minimized. In addition, previously-applied coatings do not need to be removed from the component before the rejuvenation process begins. By not removing the previously-applied coatings from the component, the present disclosure enhances the life of the part.

Referring now to the drawings,FIG. 1illustrates a schematic cross-sectional view of one embodiment of a gas turbine engine10(high-bypass type) according to the present disclosure. More specifically, the gas turbine engine10may include an aircraft engine, e.g. for an airplane, helicopter, or similar. As shown, the gas turbine engine10has an axial longitudinal centerline axis12therethrough for reference purposes. Further, as shown, the gas turbine engine10preferably includes a core gas turbine engine generally identified by numeral14and a fan section16positioned upstream thereof. The core engine14typically includes a generally tubular outer casing18that defines an annular inlet20. The outer casing18further encloses and supports a booster22for raising the pressure of the air that enters core engine14to a first pressure level. A high pressure, multi-stage, axial-flow compressor24receives pressurized air from the booster22and further increases the pressure of the air. The compressor24includes rotating blades and stationary vanes that have the function of directing and compressing air within the turbine engine10. The pressurized air flows to a combustor26, where fuel is injected into the pressurized air stream and ignited to raise the temperature and energy level of the pressurized air. The high energy combustion products flow from the combustor26to a first (high pressure) turbine28for driving the high pressure compressor24through a first (high pressure) drive shaft30, and then to a second (low pressure) turbine32for driving the booster22and the fan section16through a second (low pressure) drive shaft34that is coaxial with the first drive shaft30. After driving each of the turbines28and32, the combustion products leave the core engine14through an exhaust nozzle36to provide at least a portion of the jet propulsive thrust of the engine10.

The fan section16includes a rotatable, axial-flow fan rotor38that is surrounded by an annular fan casing40. It will be appreciated that fan casing40is supported from the core engine14by a plurality of substantially radially-extending, circumferentially-spaced outlet guide vanes42. In this way, the fan casing40encloses the fan rotor38and the fan rotor blades44. The downstream section46of the fan casing40extends over an outer portion of the core engine14to define a secondary, or bypass, airflow conduit48that provides additional jet propulsive thrust.

From a flow standpoint, it will be appreciated that an initial airflow, represented by arrow50, enters the gas turbine engine10through an inlet52to the fan casing40. The airflow passes through the fan blades44and splits into a first air flow (represented by arrow54) that moves through the conduit48and a second air flow (represented by arrow56) which enters the booster22.

The pressure of the second compressed airflow56is increased and enters the high pressure compressor24, as represented by arrow58. After mixing with fuel and being combusted in the combustor26, the combustion products60exit the combustor26and flow through the first turbine28. The combustion products60then flow through the second turbine32and exit the exhaust nozzle36to provide at least a portion of the thrust for the gas turbine engine10.

Still referring toFIG. 1, the combustor26includes an annular combustion chamber62that is coaxial with the longitudinal centerline axis12, as well as an inlet64and an outlet66. As noted above, the combustor26receives an annular stream of pressurized air from a high pressure compressor discharge outlet69. Fuel is injected from a fuel nozzle to mix with the air and form a fuel-air mixture that is provided to the combustion chamber62for combustion. Ignition of the fuel-air mixture is accomplished by a suitable igniter, and the resulting combustion gases60flow in an axial direction toward and into an annular, first stage turbine nozzle72. The nozzle72is defined by an annular flow channel that includes a plurality of radially-extending, circumferentially-spaced nozzle vanes74that turn the gases so that they flow angularly and impinge upon the first stage turbine blades43of the first turbine28. Similarly, the second stage turbine32may include a plurality of second stage turbine blades45. As shown inFIG. 1, the first turbine28preferably rotates the high-pressure compressor24via the first drive shaft30, whereas the low-pressure turbine32preferably drives the booster22and the fan rotor38via the second drive shaft34.

Referring now toFIG. 2, an exemplary turbine blade100of the gas turbine engine10ofFIG. 1is illustrated. As shown, the blade100is generally represented as being adapted for mounting to a disk or rotor (not shown) within the turbine section of the gas turbine engine10. For this reason, the turbine blade100is represented as including a dovetail102for anchoring the blade100to a turbine disk by interlocking with a complementary dovetail slot formed in the circumference of the disk. As represented inFIG. 2, the interlocking features may include protrusions referred to as tangs104that engage recesses defined by the dovetail slot. The blade100is further shown as having a platform106that separates an airfoil108from a shank105on which the dovetail102is defined. The turbine blade100also includes a blade tip109disposed opposite the platform106. As such, the blade tip109generally defines the radially outermost portion of the blade100and, thus, may be configured to be positioned adjacent to a stationary shroud (not shown) of the gas turbine engine10.

Because they are directly subjected to hot combustion gases during operation of the engine, the airfoil108, platform106, and/or blade tip109typically have very demanding material requirements. The platform106and the blade tip109are further critical regions of the turbine blade100in that they create the inner and outer flowpath surfaces for the hot gas path within the turbine section. In addition, the platform106creates a seal to prevent mixing of the hot combustion gases with lower temperature gases to which the shank105, its dovetail102, and the turbine disk are exposed. Further, the blade tip109may be subjected to creep due to high strain loads and wear interactions between it and the shroud surrounding the blade tips109. The dovetail102is also a critical region in that it is subjected to wear and high loads resulting from its engagement with a dovetail slot and the high centrifugal loading generated by the blade100.

Thus, the turbine blade100(or portions thereof) is typically coated with various coatings that are dependent on turbine operations and associated temperatures. For example, in particular embodiments, the shank105may have been previously coated with a chromium-based material, whereas the airfoil108may have been previously coated with an aluminum-based material. During operation, such coatings begin to wear off due to oxidation and other environmental conditions within the turbine engine10. Thus, the present disclosure is directed to systems and methods for rejuvenating such turbine blades. Though the present disclosure is described in reference to a turbine blade, it should be understood that the coating systems and methods as described herein may be applied to any gas turbine engine component. For example, in certain embodiments, the gas turbine engine components that may be coated according to the present disclosure in addition to the turbine blade may include but are not limited to fan blades, compressor blades, nozzles, shrouds, shroud supports, frames, turbine vanes, guide vanes, compressor vanes, or similar.

Referring now toFIG. 3, a schematic view of a kit110for rejuvenating such coated components, e.g. the coated turbine blades100, of the gas turbine engine10is illustrated. As shown, the kit110includes a first coating material112and a second, different coating material114. Although the kit110is illustrated having two coating materials, it should be understood by those of ordinary skill in the art that any number of coating materials may be used according to the present disclosure. Further, the first and second coating materials112,114may include any suitable coatings, including but not limited to chromium-based coatings, aluminum-based coatings, silicon-based coatings, platinum coatings, palladium coatings, or any other suitable coating materials. For example, the chromium-based coatings as described herein may include the chromium coating and related chromizing process as described in U.S. Patent Application Publication No.: 2010/0151124 entitled “Slurry Chromizing Process” filed on Aug. 3, 2007, which is incorporated herein by reference in its entirety. More specifically, in particular embodiments, the coating materials112,114may include aluminide (i.e. Al coating over a Ni, Co, or Fe base alloy), chromide (i.e. Cr coating over Ni, Co, or Fe base alloy), silicide (i.e. silicon coating over Ni, Co, Fe base alloy), Pt—Al (i.e. platinum plating then an aluminide coating), Pd—Al (i.e. palladium plating then an aluminide coating, Hf—Al (i.e. hafniding and aluminiding), Cr—Al on one surface with pure chromide or aluminide on another surface, Al—Si on one surface with pure Cr or Al on another surface, Al with 5% Si/Cr coating, or any other suitable coatings or combinations thereof.

The kit110also includes a maskant116for isolating a first coated portion of the component from a second coated portion of the component. More specifically, as shown in the illustrated embodiment, the maskant116may include a masking chamber118. In such an embodiment, the masking chamber118is configured to receive a first coated portion of the component, i.e. the coated shank105of the turbine blade100, so as to isolate the first coated portion from the second coated portion, i.e. the airfoil108of the turbine blade100, which is described in more detail in regards toFIG. 4below.

Further, as shown, the kit110includes a coating system120configured to simultaneously deposit the first and second coating materials112,114on the first and second coated portions105,108of the component, respectively. For example, in one embodiment, the masking chamber116may be filled with the first coating material112, e.g. a chromium-based powder. In addition, as shown, the coating system120may include a coating furnace122configured to apply the second coating material114on the second coated portion108of the component, i.e. the airfoil108of the turbine blade100as the first coating material is coating the first coated portion of the component. Thus, in such embodiments, the second coating material114may include a vapor-based aluminide material.

Referring now toFIG. 4, a schematic flow diagram of one embodiment of a process200for rejuvenating a previously coated turbine blade100of the gas turbine engine10is illustrated. As shown at202, the process200includes providing an empty masking chamber118with optional locating tangs that are configured to engage the tang104of shank105of the turbine blade100so as to retain the turbine blade100therein. Thus, as shown at204, the process200includes uninstalling the turbine blade100from the engine10and placing the coated turbine blade100within the masking chamber118. As shown at206, the process200includes filling the masking chamber118to the platform edge with the first coating material112, which as mentioned, may be a chromium-based powder. As shown at208, the process200includes placing a masking lid119on the masking chamber118so as to isolate the airfoil108of the turbine blade100from the shank105of the turbine blade100. As shown at210, the process200may also optionally include placing an additional maskant117(e.g. a maskant film or similar) around the masking lid119and the top of the masking chamber118so as to further isolate the airfoil108of the turbine blade100from the shank105of the turbine blade100. Such isolation prevents the first coating material112from damaging the airfoil108of the turbine blade100. As shown at212, the process200includes placing the turbine blade100(or a plurality of turbine blades100) in the coating furnace122. Thus, the method200includes simultaneously depositing the first coating material112on the shank105of the turbine blade100and the second coating material114on the airfoil108of the turbine blade100, i.e. via the coating furnace122.

More specifically, in certain embodiments, the coating furnace122is configured to apply the second coating material114via diffusion coating. For typical diffusion coating processes, aluminum-based materials (e.g. aluminides), chrome-based materials, and/or silicon-based materials may be mixed with an activator (e.g. a halide activator) and heated via the coating furnace122to form gaseous metal compounds which result in the deposition of the metal on the surface of the part to be coated, i.e. the airfoil108of the turbine blade100. Suitable temperatures for diffusion coating processes may be from about 1500° F. (about 815° C.) to about 2200° F. (about 1205° C.), more preferably from about 1900° F. (about 1040° C.) to about 2100° F. (about 1150° C.). Further, the turbine blade100may remain in the coating furnace122for any suitable amount of time during the diffusion coating process depending on a desired thickness of the coating. For example, in one embodiment, the turbine blade100may remain in the coating furnace from about one (1) hour to about ten (10) hours. In additional embodiments, the turbine blade100may remain in the coating furnace122for less than one hour or for greater than ten hours. Thus, the gaseous metal compounds decompose upon contact with the surfaces of the part, thereby depositing the diffusion coating on the surface thereof. In additional embodiments, the airfoil108(or second coated portion of the component) may coated using any other suitable coating techniques in addition to diffusion coating, including but not limited to slurry coating, plating, and/or pack or powder coating.

As shown at214, the process200further includes removing the turbine blade100from the coating furnace122and allowing the blade100to cool. As shown at216, the masking lid119can then be removed from atop the masking chamber118. As shown at218, the coated turbine blade100is removed from the masking chamber118. Thus, the previously-coated turbine blade100of the present disclosure is rejuvenated using a single-step coating process (i.e. both coatings are deposited onto the blade100in different locations at the same time).

Referring now toFIG. 5, a flow diagram of one embodiment of a method300for coating a component of a gas turbine engine10is illustrated. As shown at302, the method300includes uninstalling the damaged coated component from the gas turbine engine10. As mentioned, the component of the gas turbine engine10may include any suitable component such as a turbine blade, a fan blade, a compressor blade, a nozzle, shrouds, shroud supports, frames, a turbine vane, a guide vane, or a compressor vane.

As shown at304, the method300includes isolating a first coated portion of the component of the gas turbine engine from a second coated portion of the component. More specifically, where the coated component is a turbine blade100, the step of isolating the first coated portion of the blade100of the gas turbine engine10from the second coated portion of the blade100may include placing a shank of the blade100in a masking chamber and placing a masking lid on the masking chamber against a platform of an airfoil of the blade so as to isolate the shank from the airfoil.

As shown at306, the method300also includes simultaneously depositing a first coating material on the first coated portion of the component and a different, second coating material on the second coated portion of the component. In certain embodiments, the step of simultaneously depositing the first coating material on the first coated portion of the component and the second coating material on the second coated portion of the component may include diffusion coating, plating, slurry coating, powder coating, or any other suitable coating process. As shown at308, the method300includes reinstalling the rejuvenated coated component into the gas turbine engine10.

Further, the first and second coating materials may include chromium-based coatings, aluminum-based coatings, silicon-based coatings, platinum coatings, palladium coatings, or any other suitable coating materials such as those described herein. More specifically, the first coating material may include a chromium-based material, whereas the second coating material may include an aluminum-based material. In such embodiments, the step of simultaneously depositing the first coating material on the first coated portion of the component and the second coating material on the second coated portion of the component may include filling the masking chamber with the chromium-based material before placing the masking lid on the masking chamber so as to coat the shank and depositing the airfoil with the aluminum-based material.

In additional embodiments, the step of depositing the airfoil with the aluminum-based material may include placing the turbine blade in a coating furnace and exposing the airfoil to a vapor phase aluminum-based material so as to coat the airfoil, as previously described.

In yet another embodiment, the method300may further include removing the turbine blade100from the coating furnace, cooling the turbine blade100while the shank remains in the masking chamber, removing the masking lid from the masking chamber, and removing the shank from the masking chamber.