Fluid reaction surfaces (i.e. – impellers) – Specific blade structure – Having wear liner – sheathing or insert
Reexamination Certificate
2000-11-08
2004-06-29
Verdier, Christopher (Department: 3745)
Fluid reaction surfaces (i.e., impellers)
Specific blade structure
Having wear liner, sheathing or insert
C416S24100B, C415S173100, C415S173400, C428S304400, C428S306600, C428S307300, C428S307700, C428S312200
Reexamination Certificate
active
06755619
ABSTRACT:
This invention relates to aircraft gas turbines and, more particularly, to the turbine blades used in such gas turbines.
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 combustion 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 turbine blades are mounted on a turbine disk, which rotates on a shaft inside a tunnel defined by a stationary cylindrical structure termed the stationary shroud. (This “stationary shroud” is distinct from the rotating shroud found on some turbine blades.) The hot combustion gases flow from the engine's combustor and into the tunnel. The hot combustion gases pass through the turbine blade/turbine vane structure and cause the turbine to turn. To achieve a high efficiency, it is important that as little as possible of the hot combustion gases leak around the turbine by passing through the clearance between the outermost tip of the turbine blade and the innermost surface of the stationary shroud. However, the sealing of the turbine structure against such leakage presents a problem, because the components of the structure expand and contract differently during the temperature changes of over 2000° F. that are experienced during each cycle of engine operation.
To prevent the leakage of hot combustion gases around the turbine, it is known to size the components so that there is initially a very small gap between the blade tips of the turbine blades and the inner surface of the stationary shroud. As the turbine blades heat and expand when the engine is first operated, there is a contacting between the blade tips and the stationary shroud as the turbine turns. Further contacting between the blade tips and the stationary shroud sometimes occurs again during operation of the engine in the usual operating conditions or under unusual conditions such as application of emergency power or large loads applied to the components.
In these circumstances, material may be worn away from the blade tips and/or the shroud, and the clearance gap between the blade tips and the inner surface of the stationary shroud increases. Over time the efficiency of the gas turbine declines because increasing amounts of hot combustion gas leak through the enlarged clearance gap. When the efficiency has decreased by an amount that justifies a repair, the blade tips are repaired by welding new material onto the blade tips to lengthen them back to about their original lengths, and/or the shrouds are refurbished or replaced.
The repair of the blade tips is costly and difficult to perform for some configurations of the turbine blades. There is a need for an improved approach to blade tip design and materials to alleviate the problems associated with the wearing away of the blade tips.
BRIEF SUMMARY OF THE INVENTION
The present invention provides a turbine blade with a blade tip seal that is highly resistant to removal by the frictional rubbing that occurs between the blade tip seal and the inner surface of the stationary shroud. The blade tip seal has excellent high-temperature strength capability, and is operable to temperatures in excess of those encountered in current gas turbine engines. The abrasion of the blade-tip material may be controlled by selectively varying its composition. The blade tip seal is reduced in weight as compared with available blade tips, an important advantage that allows the weight of the remainder of the turbine blade, turbine disk, and supporting structure to be reduced as well. It is also more tolerant of excess-temperature excursions, so that less cooling is required and there is less concern with hot spots. The primary component of the blade-tip material is not subject to oxidation damage. The blade tip seal is resistant to the impact damage and erosion that are caused by particles that impact upon it, and it is also resistant to corrosion damage in the combustion gases.
A turbine blade comprises an airfoil section having a root end and a tip end, and an attachment at the root end of the airfoil section. A blade tip seal is joined to the tip end of the airfoil section. The blade tip seal comprises an open-cell solid ceramic foam made of ceramic cell walls having intracellular volume therebetween. The ceramic foam is open celled, so that the intracellular volume may be filled with a metal such as a nickel-base alloy or other material, or it may be porous (empty space).
In one approach, the blade tip seal comprises a blade interface adjacent to the tip end of the airfoil section, with the blade interface comprising the ceramic foam and a metal within the intracellular volume. The metal within the intracellular volume in this blade interface region aids in bonding the blade tip seal to the airfoil section of the metallic turbine blade. Such a blade tip seal may further comprise a contact region remote from the blade interface, with the contact region comprising the ceramic foam and a porous (empty) intracellular volume. The contact region is thus a ceramic skeleton that is self supporting but porous. Although porous, the contact region still serves to prevent combustion gas leakage therethrough. It has sufficient strength, is abrasive, and is impact resistant. The ceramic skeleton is preferably aluminum oxide, but its abrasive properties may be varied with compositional changes such as the addition of an abrasive ceramic mixed with the base ceramic.
The blade tip seal desirably comprises an amount exceeding about 60 volume percent of ceramic, and preferably comprises from about 60 to about 80 percent by volume of ceramic. The physical form of the blade tip seal may be varied as necessary, such as a plate form or a rim form.
The blade tip seal is preferably prepared separately from the airfoil section of the turbine blade, and then the blade tip seal and the airfoil are joined together by an appropriate process such as welding or diffusion bonding. The blade tip seal may be prepared by providing a piece of a sacrificial ceramic having the shape of the blade tip seal, and contacting the piece of the sacrificial ceramic with a reactive metal which reacts with the sacrificial ceramic to form an oxidized ceramic of the reactive metal and a reduced form of the ceramic. The resulting structure comprises an open-celled ceramic foam of the oxidized ceramic compound of the reactive metal with ceramic cell walls and the intracellular volume between the ceramic cell walls. The contact region is formed by removing metal from the contact region, leaving the ceramic foam of the contact region with its intracellular volume as porosity. The metal may be removed by any appropriate technique, such as chemically attacking the metal in the contact region during the fabrication processing, or thermally attacking the metal in the contact region during operation of the engine.
The present approach provides a blade tip seal that has sufficient strength and is also abrasive so that it wears away the stationary shroud material, rather than being worn away by the stationary shroud material. The wearing away of the stationary shroud material is preferred to the wearing away of the blade tip to promote roundness of the shroud assembly. Otherwise, one mispositioned shroud could cause all blades to become shorter, increasing leakage around the entire circumference of the turbine. The reduced weight at the furthest extent of the turbine blade, which rotates rapidly on the turbine shaft, allows the reduction in material and weight in other components, such as the remainder of the turbine blade, the turbine disk, the turbine shaft, bearings and other related structure, and the structure that supports the gas turbine engine.
REFERENCES:
patent: 437
Austin Curtiss Mitchell
Grylls Richard John
Schell Jerry Donald
Garmong Gregory O.
General Electric Company
Maria Carmen Santa
McNees Wallace & Nurick LLC
Verdier Christopher
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