Fluid reaction surfaces (i.e. – impellers) – Specific blade structure – Coating – specific composition or characteristic
Reexamination Certificate
1999-09-24
2001-09-04
Lopez, F. Daniel (Department: 3745)
Fluid reaction surfaces (i.e., impellers)
Specific blade structure
Coating, specific composition or characteristic
C416S24100B
Reexamination Certificate
active
06283715
ABSTRACT:
FIELD OF THE INVENTION
This invention relates to aircraft gas turbine engines, and, more particularly, to protective coatings placed on turbine components such as turbine blades and turbine vanes.
BACKGROUND OF THE INVENTION
In an aircraft gas turbine (jet) engine, air is drawn into the front of the engine, compressed by a shaft-mounted compressor, and mixed with fuel. The mixture is burned, and the hot exhaust gases are passed through a turbine mounted on the same shaft. The flow of combustion gas turns the turbine by impingement against an airfoil section of the turbine blades and vanes, which turns the shaft and provides power to the compressor and fan. The hot exhaust gases flow from the back of the engine, driving it and the aircraft forwardly.
The hotter the combustion and exhaust gases, the more efficient is the operation of the jet engine. There is thus an incentive to raise the combustion and exhaust gas temperatures. The maximum temperature of the combustion gases is normally limited by the materials used to fabricate the turbine vanes and turbine blades of the turbine, upon which the hot combustion gases impinge. In current engines, the turbine vanes and blades are made of nickel-based superalloys, and can operate at temperatures of up to about 1900-2100° F.
Many approaches have been used to increase the operating temperature limit of the turbine blades and vanes to their current levels. For example, the composition and processing of the base materials themselves have been improved.
Physical cooling techniques may also be used. In one technique, internal cooling passages through the interior of the turbine airfoil are present. Air is forced through the cooling passages and out openings at the external surface of the airfoil, removing heat from the interior of the airfoil and, in some cases, providing a boundary layer of cooler air at the surface of the airfoil. In another approach, a protective layer or a metal/ceramic thermal barrier coating (TBC) system is applied to the airfoil, which acts as a substrate.
The gas turbine blade or vane is operated in a highly aggressive environment that can cause damage to the component in service. The environmental damage may be in various forms, such as particle erosion, different types of corrosion, and oxidation, and complex combinations of these damage modes, in the hot combustion gas environment. The rate of environmental damage may be lessened somewhat with the use of the protective layers. However, the various types of environmental damage are still observed, often necessitating premature replacement or repair of components after service exposure.
There is a need for an improved approach to the protection of gas turbine components such as turbine blades and vanes. The present invention fulfills this need, and further provides related advantages.
BRIEF SUMMARY OF THE INVENTION
The present invention provides a coated gas turbine component and a method for its preparation. The coating achieves protection of the component in regions usually subject in service to intermediate-temperature corrosion. The technique of the invention is fully compatible with other coating procedures, such as the application of a thermal barrier coating.
A coated gas turbine component comprises a gas turbine component formed of a base metal. The gas turbine component includes a platform, a shank extending downwardly from the platform, and an airfoil extending upwardly from the platform. A first coating contacts the base metal of at least a portion of the shank and is interdiffused therewith. The first coating comprises a first chromide layer. A second coating contacts at least a portion of the airfoil. The second coating comprises a protective layer that is a diffusion coating or an overlay coating.
Preferably but not necessarily, a second chromide layer is disposed between the base metal of the airfoil and the protective layer, and is interdiffused with the base metal of the airfoil. Where there is no second chromide layer, the protective layer of the second coating directly contacts the base metal of the airfoil. A ceramic layer may overlie the protective layer of the second coating.
The base metal is preferably a nickel-base superalloy, and may be of any operable type and composition. The first chromide layer preferably comprises from about 20 to about 30 weight percent chromium, and is from about 0.001 to about 0.002 inch thick. The second chromide layer, where present, also preferably comprises from about 20 to about 30 weight percent chromium, and is from about 0.001 to about 0.002 inch thick. The first and second chromide layers further optionally include chromide-modifying elements such as silicon deposited with the chromium in forming the chromide layer, elements interdiffused with the chromide layer from the base metal, and impurities. Where the chromide layer lies between the base metal and a diffusion coating, the chromide layer further includes elements interdiffused from the diffusion coating. The diffusion coating of the second coating preferably is a diffusion aluminide or a diffusion platinum aluminide, while the overlay coating is preferably an MCrAIX coating. In addition to the regions coated as discussed above, the first coating may be applied to the bottom side of the platform and some portions of the dovetail, and the second coating may be applied to the top side of the platform.
A method for preparing a coated gas turbine component comprising a base metal includes the steps of furnishing a gas turbine component comprising a platform, a shank extending downwardly from the platform, and an airfoil extending upwardly from the platform, and first applying a first coating overlying the base metal of at least a portion of the shank and interdiffused therewith. The first coating comprises a first chromide layer. The method further includes second applying a second coating overlying at least a portion of the airfoil. The second coating comprises a protective layer selected from the group consisting of a diffusion coating and an overlay coating. Optionally, the step of second applying includes steps of depositing a second chromide layer overlying the base metal of the airfoil, and depositing the protective layer overlying the second chromide layer. Where no second chromide layer is used, the protective layer is deposited directly overlying and contacting the base metal of the airfoil.
The first chromide layer is formed by depositing chromium overlying the base metal of the shank, and interdiffusing the chromium with the base metal. The second chromide layer, where present, is deposited by depositing chromium overlying the base metal of the airfoil, and interdiffusing the chromium with the base metal. In either case, the chromide preferably comprises about 20 to about 30 weight percent chromium. Where both first and second chromide layers are present, they are preferably deposited concurrently. After diffusion, the two chromide coatings may have substantially the same composition, or they may have different compositions. The step of second applying may optionally include applying a ceramic thermal barrier coating overlying the protective layer of the second coating.
The present approach results in a chromide coating on the shank, and optionally the underside of the platform and portions of the dovetail. The chromide coating protects the shank against intermediate-temperature corrosion in the range of about 1100° F.-1500° F., the operating temperature of the shank in most service environments. The airfoil is optionally coated with chromide coating, and then further coated with the protective layer, for enhanced corrosion resistance. All or portions of the airfoil that are subjected to the highest temperatures, such as the concave high-pressure side, may be further coated with the ceramic thermal barrier coating.
Other features and advantages of the present invention will be apparent from the following more detailed description of the preferred embodiment, taken in conjunction with the accompanying drawings, which illustrate, by way of ex
Estill Eric A.
Lee Elissa H.
Maurer Norbert O.
Nagaraj Bangalore A.
General Electric Company
Hess Andrew C.
Lopez F. Daniel
Narcisco David L.
Nguyen Ninh
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