Fluid reaction surfaces (i.e. – impellers) – Specific blade structure – Coating – specific composition or characteristic
Reexamination Certificate
2002-05-06
2004-02-24
Nguyen, Ninh (Department: 3745)
Fluid reaction surfaces (i.e., impellers)
Specific blade structure
Coating, specific composition or characteristic
Reexamination Certificate
active
06695587
ABSTRACT:
This invention relates to protective coatings on articles, and, more particularly, to aluminide and platinum-aluminide coatings on aircraft gas turbine components having airfoils.
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 combusted, and the resulting hot combustion gases are passed through a turbine mounted on the same shaft. The flow of gas turns the turbine by contacting an airfoil portion of the turbine blade, which turns the shaft and provides power to the compressor. The hot exhaust gases flow from the back of the engine, driving it and the aircraft forwardly.
The hotter the turbine gases, the more efficient is the operation of the jet engine. There is thus an incentive to raise the turbine operating temperature. However, the maximum temperature of the turbine gases is normally limited by the materials used to fabricate the turbine vanes and turbine blades of the turbine. In current engines, the turbine vanes and blades are made of nickel-based or cobalt-based superalloys that can operate at temperatures of up to 1900-2100° F.
Many approaches have been used to increase the operating temperature limits and operating lives of the airfoils of the turbine blades and vanes. The compositions and processing of the materials themselves have been improved. The articles may be formed as oriented single crystals to take advantage of superior properties observed in certain crystallographic directions. Physical cooling techniques are used. For example, internal cooling channels may be provided within the components, and cooler air is forced through the channels during engine operation.
In another approach, a protective layer is applied to the airfoil of the turbine blade or turbine vane component, which acts as a substrate. Among the currently known diffusional protective layers are aluminide and platinum aluminide layers. The protective layer protects the substrate against environmental damage from the hot, highly corrosive combustion gases. This protective layer, with no overlying ceramic layer, is useful in intermediate-temperature applications. For higher temperature applications, a ceramic thermal barrier coating layer may be applied overlying the protective layer, to form a thermal barrier coating (TBC) system. The ceramic thermal barrier coating layer insulates the component from the exhaust gas, permitting the exhaust gas to be hotter than would otherwise be possible with the particular material and fabrication process of the substrate.
Even with the use of these protective techniques, there remain problems to overcome in extending the operating service temperatures and operating lives of the turbine blade components. For example, some portions of the airfoil have been observed to fail prematurely due to low-cycle fatigue, wherein that portion of the airfoil is subjected to repetitive, relatively large strain cycles at elevated temperature. There is a need for an approach to overcoming such problems, while retaining the benefits of the available protection techniques.
BRIEF SUMMARY OF THE INVENTION
The present invention provides a technique for reducing the susceptibility of gas turbine components to property degradation such as low-cycle fatigue failures, while retaining the benefits associated with protective coatings that are applied to the components. The present approach takes a highly selective approach to the protection of the turbine components, optimizing the performance of the protective system at various portions of the component. Expensive platinum is conserved, although this is a relatively minor benefit. The present approach may be accomplished as part of the normal production operation, without major modifications.
A method for preparing an article protected by an aluminide coating is utilized with a component article of a gas turbine having an airfoil section and comprising a nickel-base superalloy. The method includes masking a portion of the airfoil section, leaving an unmasked portion of the airfoil section, and depositing a noble metal such as platinum onto the airfoil section as a substrate. The result is that the unmasked portion has the noble metal layer thereon and the masked portion has no noble metal layer thereon. The mask is removed, and an aluminum-containing layer is deposited onto the airfoil section of the article. Typically the noble metal, the aluminum-containing layer, and the substrate material are interdiffused. A ceramic layer may be deposited over the aluminum-containing layer, to form a thermal barrier coating.
In the preferred application of the present invention, the selective use of the noble metal allows a reduction in premature failures due to low-cycle fatigue. The application is based upon the recognition that platinum enhances the attraction of aluminum to the coating. For a part coated in an atmosphere of constant aluminum activity, the area with platinum will coat thicker and have a higher total aluminum content than an area without platinum. Thick coatings are more prone to mechanical property degradation such as low-cycle-fatigue cracking. The incidence of low-cycle fatigue damage may be lessened in some areas of the airfoil by using an aluminide protective coating rather than a platinum aluminide protective coating.
Consistent with this approach, an area of the article that is subject to mechanical property degradation such as low-cycle fatigue damage is identified. The masked portion includes the area of the article that is subject to such mechanical property degradation in the form of low-cycle fatigue damage. A region of particular concern is the portion of the airfoil adjacent to a trailing edge of the airfoil, and most particularly the trailing edge adjacent to a platform portion of the turbine component. The trailing edge region adjacent to the platform experiences less severe temperatures than the leading edge-region, so the use of the aluminide coating at the trailing edge root is sufficient from a protection standpoint.
The result is an article protected by an aluminide coating, comprising a nickel-base superalloy substrate in the form of a gas turbine component article having an airfoil section, a platinum aluminide coating in an unmasked portion of the airfoil, and an aluminide coating in a masked portion and in the unmasked portion of the airfoil. The masked portion is the portion of the airfoil that is identified as most susceptible to mechanical property degradation in the form of low-cycle fatigue damage. In the case of most interest, this portion is the trailing edge root of the airfoil, as discussed above.
The present approach thus provides a technique for selectively coating the gas turbine component to protect the component yet to reduce the incidence of property degradation such as low-cycle fatigue failures. 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 example, the principles of the invention. The scope of the invention is not, however, limited to this preferred embodiment.
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Clarke Jonathan P.
Maricocchi Antonio F.
Wustman Roger D.
Garmong Gregory O.
Narciso David L.
Nguyen Ninh
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