Fluid reaction surfaces (i.e. – impellers) – Specific blade structure – Coating – specific composition or characteristic
Reexamination Certificate
1999-08-11
2001-09-04
Look, Edward K. (Department: 3745)
Fluid reaction surfaces (i.e., impellers)
Specific blade structure
Coating, specific composition or characteristic
C416S24100B, C415S217100
Reexamination Certificate
active
06283714
ABSTRACT:
FIELD OF THE INVENTION
This invention relates to a gas turbine airfoil having an internal cooling passage, and, more particularly, to the protection of the internal and external surfaces of such a gas turbine airfoil.
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 metal 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. 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.
In another approach, a protective layer or a ceramic/metal thermal barrier coating (TBC) system is applied to the airfoil, which acts as a substrate. The protective layer with no overlying ceramic layer (in which case the protective layer is termed an “environmental coating”) is useful in intermediate-temperature applications. The currently known protective layers include diffusion aluminides and overlays. A ceramic thermal barrier coating layer may be applied overlying the protective layer on the external airfoil surface, to form a thermal barrier coating system (in which case the protective layer is termed a “bond coat”). The thermal barrier coating system is useful in higher-temperature applications. The ceramic thermal barrier coating insulates the component from the combustion gas, permitting the combustion gas to be hotter than would otherwise be possible with the particular material and fabrication process of the substrate.
During normal service of a gas turbine blade or vane, the airfoil layer is typically damaged by particle impacts and by oxidation/corrosion in the hot combustion gas environment. If the damage is not too severe, the gas turbine blade or vane may be removed from service, repaired, and returned to service. The repair typically includes, among other things, stripping away the damaged protective coating and thermal barrier coating layer, if any, from the external airfoil surface, and applying new protective coatings.
For those cases where the airfoil has internal cooling passages, the removal of the external protective coating during repair operations reduces the remaining structural thickness of base metal that lies between the internal cooling passage and the external airfoil surface. The inventors have determined that this is particularly a concern when the external protective coating, or an inner portion of the external protective coating, experiences significant diffusion into the base metal either during manufacturing or during service. For example, a typical 0.002 inch thick diffusion aluminide coating includes a diffusion zone that is about 0.001 inch thick, and an “add-on” layer that is about 0.001 inch thick. The diffusion zone consumes a portion of the wall of the airfoil, reducing its effective thickness for supporting loads. When this thickness becomes so reduced that it can no longer support the required structural loads, the turbine blade or vane becomes unrepairable, and must be discarded even though otherwise it could be repaired and returned to service.
There is a need for an improved approach to the protection of gas turbine airfoils containing internal cooling passages, which permits their repeated repair without loss of structural integrity. The present invention fulfills this need, and further provides related advantages.
BRIEF SUMMARY OF THE INVENTION
The present approach provides a technique for protecting the internal cooling passage walls and also the external surfaces of a gas turbine airfoil. The technique allows the wall thickness between the internal cooling passages and the external surface to be maintained after one or more repair operations in which the external protective coating is removed and replaced. Airfoils which previously would have been unrepairable can now be repaired with confidence. Designers of airfoils can therefore select wall thicknesses with the knowledge that they will be maintained even after repairs, allowing the wall thicknesses to be selected for maximum cooling efficiency and structural integrity, and reduced rotor weight. The result is an improved thrust-to-weight ratio of the gas turbine engine.
A gas turbine component includes a gas turbine airfoil formed of a base metal and having an internal cooling passage therein defined by an internal airfoil surface, and an external airfoil surface. A method for preparing a coated gas turbine airfoil comprises the step of forming a diffusion aluminide protective layer at the internal airfoil surface of the internal cooling passage. There is substantially no aluminum deposited on the external airfoil surface of the gas turbine airfoil during the internal diffusion aluminiding. The method further includes depositing an overlay protective coating on the external airfoil surface of the gas turbine airfoil, there being substantially no diffusion aluminide between the overlay protective coating and the external airfoil surface.
The diffusion aluminide protective layer at the internal airfoil surface of the internal cooling passage is preferably formed by contacting aluminum-containing compounds to the internal airfoil surface to deposit aluminum thereon, and interdiffusing the aluminum into the base metal. Modifying elements such as hafnium, zirconium, yttrium, silicon, titanium, tantalum, cobalt, chromium, platinum, and palladium, and combinations thereof may be applied with the aluminum, and diffused into the base metal with the aluminum. The aluminum and optional modifying elements are applied by any operable technique, such as slurry coating, foam coating, chemical vapor deposition, organo-metallic chemical vapor deposition, pack cementation, or vapor phase aluminiding. When the airfoil is a turbine blade, the aluminum is preferably flowed from a root end of the internal cooling passage toward a tip end of the internal cooling passage. In all cases, care is taken that aluminum does not deposit onto the external surface of the airfoil.
The overlay protective coating is deposited on the external surface of the airfoil, preferably as an MCrAlX alloy. A ceramic coating of, for example, yttria-stabilized zirconia may be deposited overlying the overlay protective coating. The overlay protective coating and the ceramic coating are preferably applied by a technique such as thermal spray, air plasma spray, low pressure plasma spray, high-velocity oxyfuel, ion plasma deposition, electron beam physical vapor deposition, sputtering, c
Darolia Ramgopal
Rigney Joseph D.
Weimer Michael J.
General Electric Company
Hess Andrew C.
Look Edward K.
Narciso David L.
Nguyen Ninh
LandOfFree
Protection of internal and external surfaces of gas turbine... does not yet have a rating. At this time, there are no reviews or comments for this patent.
If you have personal experience with Protection of internal and external surfaces of gas turbine..., we encourage you to share that experience with our LandOfFree.com community. Your opinion is very important and Protection of internal and external surfaces of gas turbine... will most certainly appreciate the feedback.
Profile ID: LFUS-PAI-O-2445821