Stock material or miscellaneous articles – Composite – Of inorganic material
Reexamination Certificate
2000-11-06
2003-01-28
Jones, Deborah (Department: 1775)
Stock material or miscellaneous articles
Composite
Of inorganic material
C428S632000, C428S469000, C428S218000, C428S304400, C428S307300, C428S309900, C428S699000, C428S702000, C428S131000, C428S133000, C428S140000, C416S24100B, C416S09700R
Reexamination Certificate
active
06511762
ABSTRACT:
FIELD OF THE INVENTION
This invention relates generally to gas turbine engines, and in particular, to a cooled flow path surface region.
BACKGROUND OF THE INVENTION
In gas turbine engines, for example, aircraft engines, air is drawn into the front of the engine, compressed by a shaft-mounted rotary-type compressor, and mixed with fuel. The mixture is burned, and the hot exhaust gases are passed through a turbine mounted on a shaft. The flow of gas turns the turbine, which turns the shaft and drives the compressor and fan. The hot exhaust gases flow from the back of the engine, driving it and the aircraft forward.
During operation of gas turbine engines, the temperatures of combustion gases may exceed 3,000° F., considerably higher than the melting temperatures of the metal parts of the engine, which are in contact with these gases. Operation of these engines at gas temperatures that are above the metal part melting temperatures is a well established art, and depends in part on supplying a cooling air to the outer surfaces of the metal parts through various methods. The metal parts of these engines that are particularly subject to high temperatures, and thus require particular attention with respect to cooling, are metal parts forming combustors and parts located aft of the combustor including turbine blades, turbine vanes and exhaust nozzles.
The hotter the turbine inlet gases, the more efficient is the operation of the jet engine. There is thus an incentive to raise the turbine inlet gas temperature. However, the maximum temperature of the turbine inlet 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 superalloys, and can operate at metal surface temperatures of up to 2100°-2200° F.
The metal temperatures can be maintained below melting levels with current cooling techniques by using a combination of improved cooling designs and insulating thermal barrier coatings (TBCs). For example, with regard to the metal blades and vanes employed in aircraft engines, some cooling is achieved through convection by providing passages for flow of cooling air internally within the blades so that heat may be removed from the metal structure of the blade by the cooling air. Such blades essentially have intricate serpentine passageways within structural metal forming the cooling circuits of the blade.
Small internal orifices have also been devised to direct this circulating cooling air directly against certain inner surfaces of the airfoil to obtain cooling of the inner surface by impingement of the cooling air against the surface, a process known as impingement cooling. In addition, an array of small holes extending from the hollow core through the blade shell can provide for bleeding cooling air through the blade shell to the outer surface where a film of such air can protect the blade from direct contact with the hot gases passing through the engines, a process known as film cooling.
In another approach, a TBC is applied to the turbine blade component, which forms an interface between the metallic component and the hot gases of combustion. The TBC includes a ceramic coating that is applied to the external surface of metal parts within engines to impede the transfer of heat from hot combustion gases to the metal parts, thus insulating the component from the hot combustion gas. This permits the combustion gas to be hotter than would otherwise be possible with the particular material and fabrication process of the component. TBCs have also been used in combination with film cooling techniques wherein an array of fine holes extends from the hollow core through the TBC to provide cooling air onto the outer surface of the TBC.
TBCs are well-known ceramic coatings, for example, yttrium-stabilized zirconia (YSZ). Ceramic TBCs usually do not adhere well directly to the superalloys used in the substrates. Therefore, an additional metallic layer called a bond coat is placed between the substrate and the thermal barrier coating. The bond coat may be made of a nickel-containing overlay alloy, such as a NiCrAlY or a NiCoCrAIY, or other composition more resistant to environmental damage than the substrate, or alternatively, the bond coat may be a diffusion nickel aluminide or platinum aluminide, whose surface oxidizes to a protective aluminum oxide scale that provides improved adherence to the ceramic top coatings. The bond coat and the overlying TBC are frequently referred to as a thermal barrier coating system.
Multi layer coatings are known in the art. For example, U.S. Pat. No. 5,846,605 to Rickerby, et al., is directed to a coating having a plurality of alternate layers having different structures that produce a plurality of interfaces. The interfaces provide paths of increased resistance to heat transfer to reduce thermal conductivity. A bond coat overlying a metallic substrate is bonded to a TBC. The TBC comprises a plurality of layers, each layer having columnar grains, the columnar grains in each layer extending substantially perpendicular to the interface between the bond coating and metallic substrate. The structure is columnar to ensure that the strain tolerance of the ceramic TBC is not impaired. The difference in structure of the layers is the result of variations in the microstructure and/or density/coarseness of the columnar grains of the ceramic.
U.S. Pat. No. 5,705,231 to Nissley et al. is directed to a segmented abradable ceramic coating system having enhanced abradability and erosion resistance. A segmented abradable ceramic coating is applied to a bond coat comprising three ceramic layers that are individually applied. There is a base coat foundation layer, a graded interlayer, and an abradable top layer. The coating is characterized by a plurality of vertical microcracks.
U.S. Pat. No. 4,503,130 to Bosshart et al. is directed to coatings having a low stress to strength ratio across the depth of the coating. Graded layers of metal/ceramic material having increasing ceramic composition are sequentially applied to the metal substrate under conditions of varied substrate temperature.
U.S. Pat. No. 6,045,928 to Tsantrizos et al. is directed to a TBC comprising an MCrAlY bond coat and a dual constituent ceramic topcoat. The topcoat comprises a monolithic zirconia layer adjacent to the bond coat, a monolithic layer of calciasilica representing the outer surface of the TBC and a graded interface between the two to achieve good adhesion between the two constituents to achieve an increased thickness of the topcoat, thereby, providing for a greater temperature drop across the TBC system.
U.S. Pat. No. 4,576,874 to Spengler et al. is directed to a coating to increase resistance to spalling and corrosion. The coating is not intended to be a thermal barrier coating. A porous ceramic is applied over a MCrAlY bond coat and a dense ceramic is then applied over the porous ceramic. The porous portion is a transition zone to allow for differences in thermal expansion and provides little thermal insulation.
Improved environmental resistance to destructive oxidation and hot corrosion is desirable. In some instances, the alloying elements of the bond coat can interdiffuse with the substrate alloy and consume walls of the turbine airfoils, i.e., reduce load carrying capability. This interdiffusion also reduces environmental resistance of the coating. Even with the use of advanced cooling designs and thermal barrier coatings, it is also desirable to decrease the requirement for cooling; because reducing the demand for cooling is also well known to improve overall engine operating efficiency.
While superalloys coated with thermal barrier coating systems do provide substantially improved performance over uncoated materials, there remains room for improvement. Film cooling is achieved by passing cooling air through discrete film cooling holes, typically ranging from 0.015″ to about 0.030″ in hole diameters. The film cooling holes are typically drilled with laser
Darolia Ramgopal
Lee Ching-Pang
Schafrik Robert Edward
General Electric Company
Jones Deborah
McNees Wallace & Nurick
McNeil Jennifer
Narciso David L.
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