Rotary kinetic fluid motors or pumps – Working fluid passage or distributing means associated with... – Specific casing or vane material
Reexamination Certificate
1999-08-11
2001-08-14
Verdier, Christopher (Department: 3745)
Rotary kinetic fluid motors or pumps
Working fluid passage or distributing means associated with...
Specific casing or vane material
C416S22900R, C416S24100B, C416S24100B, C029S889700, C427S181000, C427S182000, C427S454000, C427S237000, C427S252000, C428S469000, C428S472000, C428S632000
Reexamination Certificate
active
06273678
ABSTRACT:
FIELD OF THE INVENTION
This invention relates to a gas turbine component having an internal cooling passage, and, more particularly, to the protection of the surface of the internal passage of such a gas turbine component.
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. 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 1800-2100° 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 are located within the interior of the turbine component. 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. To attain maximum cooling efficiency, the cooling passages are placed as closely to the external surface of the airfoil as is consistent with maintaining the required mechanical properties of the airfoil, to as little as about 0.020 inch in some cases.
The surfaces of the internal cooling passages and the external surfaces of the turbine component may be protected with a protective coating. Aluminide diffusion coatings are used for the internal surfaces, and aluminide diffusion coatings or overlay coatings are used on the external surfaces. A ceramic layer may also overlie the protective coating on the external surfaces. Although these internal and external protective layers provide improved resistance to environmental damage of the turbine component and the ability to operate at higher temperatures, there is an opportunity for improvement. Thus, there is a need for improved protective coating systems that extend the capabilities of the turbine components even further. The present invention fulfills this need, and further provides related advantages.
BRIEF SUMMARY OF THE INVENTION
The present invention provides an article and a method for its preparation. The article is preferably a component of a gas turbine having internal passages therein, such as the passages that channel a flow of cooling air through the component. The present approach increases the environmental resistance of the internal surfaces that form the internal passages, thereby increasing their durability. The improved environmental resistance is achieved using a modification of an existing, well-proved technology. The present approach is specific to the protection of internal surfaces, but it may be utilized in conjunction with any approach for protecting the external surfaces.
An article comprises a gas turbine component having a substrate, an internal passage through the substrate defining an internal surface of the substrate, and an external surface of the substrate. An internal protective layer overlies the internal surface of the substrate. The internal protective layer has a composition comprising aluminum, plus, in weight percent, on average from about 0.1 to about 5.0 percent of a modifying element including hafnium, yttrium, zirconium, chromium, and/or silicon, and combinations thereof.
A related method for preparing an article comprises the steps of providing a substrate having an internal passage therethrough defining an internal surface of the substrate, depositing a layer onto the internal surface comprising aluminum and a modifying element selected from the group consisting of hafnium, yttrium, zirconium, chromium, and silicon, and combinations thereof, and heating the layer comprising aluminum and the modifying element so that the aluminum and the modifying element diffuse into the substrate. The diffused material forms an internal protective layer having an average composition of from about 16 to about 30 weight percent aluminum, from about 0.1 to about 5.0 weight percent of the modifying element, and other elements interdiffused from the substrate.
The gas turbine component is preferably a gas turbine blade or gas turbine vane, with internal cooling passages. Such an article is typically made of a nickel-base superalloy. In most cases, an external protective layer in the form of a diffusion aluminide or an overlay coating is also used, optionally with the application of a ceramic layer to form a thermal barrier coating.
The present invention is used solely in conjunction with the internal surfaces of the gas turbine component and to protect these internal surfaces. The protection of the internal surfaces poses a substantially different problem than the protection of the external surfaces of the gas turbine component. The internal surfaces are usually formed by small internal passages, that are typically from about 0.1 inch to about 0.5 inch in diameter. The internal surfaces are not accessible to many types of coating techniques, such as those employing line-of-sight deposition processes. The protective layer on the internal surfaces cannot be readily repaired, and therefore must last longer than the protective layer on the external surfaces, which can be refurbished. Additionally, the internal surfaces are subjected to a significantly different service environment than the external surfaces. The external surfaces experience hot corrosion, hot oxidation, and erosion in the combustion gas. On the other hand, a flow of bleed air from the engine compressor, not combustion gas, is passed through the internal passages, and the internal surfaces are at a lower temperature than the external surfaces. The bleed cooling air typically contains salt, sulfur, and other corrodants drawn into the compressor of the engine. The presence of the combination of salt and sulfur at a temperature in the range of about 1300° F., a typical temperature for the internal surfaces, may lead to severe Type II hot corrosion on the internal surfaces. The internal surfaces of the internal passages are additionally subjected to low-to-medium temperature oxidation. The internal surfaces of the gas turbine components are thus subjected to environmental damage of a type substantially different from that experienced on the external surfaces.
The present approach provides an internal protective layer tailored to the requirements of the internal surfaces of the gas turbine component. 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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General Electric Company
Hess Andrew C.
Narciso David L.
Verdier Christopher
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