Fluid reaction surfaces (i.e. – impellers) – Specific blade structure – Coating – specific composition or characteristic
Reexamination Certificate
1999-08-11
2001-10-02
Look, Edward K. (Department: 3745)
Fluid reaction surfaces (i.e., impellers)
Specific blade structure
Coating, specific composition or characteristic
C415S217100
Reexamination Certificate
active
06296447
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
This invention provides a gas turbine component that is protected against environmental damage in different locations by different types of protective layers and coatings. The invention recognizes that different regions of the surfaces of the turbine components experience different types of degradation due to the environment, even though the different regions may be separated by a matter of inches or less. The understanding of the performance of different protective layers has progressed to the point that various protective layers may be optimized for performance under these different conditions of environmental damage. However, at this point no one type of protective layer is optimal for use in all or even a wide range of situations. The protective layer for use in each location is selected according to its performance in the service environment experienced at each location and the ability to apply the protective layers in those regions.
A gas turbine component comprises a platform having a first coating on at least a portion thereof, and an airfoil extending from the platform. The airfoil has an external surface and an internal surface defining cooling passages therethrough. The airfoil comprises an airfoil shape comprising a substrate alloy, a second coating on the external surface of the airfoil shape, the second coating being different in composition from the first coating, and a third coating on the internal surface of the airfoil shape, the third coating being different in composition from the second coating, and, preferably, from the first coating as well. The gas turbine component may also have a tip cap remote from the platform with the same composition as the substrate alloy, or fabricated primarily with another oxidation-resistant nickel-base superalloy. A fourth coating may overlie the tip cap.
At least a portion of the platform, preferably its underside remote from the airfoil, is coated with the first coating. As used herein, “coating” encompasses all operable types of protective layers, including for example diffusion coatings, overlay coatings, and thermal barrier coatings using bond coats. The platform runs cooler than the airfoil portion of the component and is exposed to less than the full velocity of the hot gas stream. Corrosion of the platform is a greater concern than oxidation and erosion. The first coating is therefore preferably optimized for corrosion resistance. The first coating is preferably a modified diffusion aluminide comprising aluminum and an element selected from the group consisting of chromium, hafnium, silicon, zirconium, yttrium, platinum, and palladium, and combinations thereof, first applied to the surface of the substrate and then interdiffused with the substrate alloy.
At least a portion of the external surface of the airfoil is coated with the second coating that is preferably optimized for performance in an oxidation and erosion environment, and, in the case of thermal barrier coating applications, additionally optimized for thermal insulation. The second coating is preferably an MCrAlX overlay coating. M refers to nickel, cobalt, iron, and combinations thereof. Cr refers to chromium, although in some of these protective coatings the chromium may be omitted. Al refers to aluminum. The X denotes elements such as hafnium, zirconium, yttrium, tantalum, rhenium, platinum, silicon, titanium, boron, carbon, and combinations thereof. The second coating may instead be a diffusion aluminide, to which is added noble metals, rare earth elements, chromium, and/or other elements or phases which increase the strength of the second coating. In either case, a ceramic layer is optionally applied overlying the metallic or intermetallic coating to form a thermal barrier coating (TBC).
Different regions of the external surface of the airfoil may be coated with different coatings. The second coating just described is typically used on a concave pressure side of the airfoil, against which the hot combustion gas directly impinges. A different coating may be used on a convex suction side (i.e., back side) of the airfoil. This portion of the airfoil operates at nearly the same temperature as the pressure side, but is less affected by hot erosion. The substrate alloy still requires protection from the environment, and specifically oxidation and hot corrosion caused by the hot combustion gas. Aerodynamic and fluid-dynamic considerations indicate that the corrosive deposits are less likely to deposit on the convex suction side of the airfoil and are more likely to stick to the concave pressure side of the airfoil. As a result, the protective coating applied on the surface of the suction side of the airfoil need not be as environmentally resistant as the protective coating on the pressure side. For example, a diffusion aluminide similar to that described above as the first coating may be used on the surface of the suction side of the airfoil.
The internal surfaces of the airfoil, defined by internal cooling passages through the airfoil, are coated with an oxidation-resistant third coating material. The demands on this third coating are less severe than those made on the other coatings, because the interior surfaces do not experienc
Nagaraj Bangalore A.
Pfaendtner Jeffrey A.
Rigney Joseph D.
General Electric Company
Hess Andrew C.
Look Edward K.
Narciso David L.
Nguyen Ninh
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