Abstract: A gas turbine engine turbine blade comprising an airfoil section, a platform section, an under platform section, and a dovetail section, an exterior surface of the dovetail section comprising a shank exterior surface and a serrated exterior surface. The blade further comprises a silicon-modified diffusion aluminide layer a surface of a turbine blade section selected from the group consisting of an exterior surface of the under platform section, the exterior surface of the dovetail section, and combinations thereof, the silicon modified diffusion aluminide layer having a concentration of silicon at a surface of the silicon-modified diffusion aluminide layer in the range of about 1 weight percent to about 10 weight percent and a concentration of aluminum at the surface of the silicon modified diffusion aluminide layer in the range of about 5 weight percent to about 25 weight percent.

Abstract: A gas turbine engine turbine blade comprising an airfoil section, a platform section, an under platform section, and a dovetail section, an exterior surface of the dovetail section comprising a shank exterior surface and a serrated exterior surface. The blade further comprises a silicon-modified diffusion aluminide layer a surface of a turbine blade section selected from the group consisting of an exterior surface of the under platform section, the exterior surface of the dovetail section, and combinations thereof, the silicon modified diffusion aluminide layer having a concentration of silicon at a surface of the silicon-modified diffusion aluminide layer in the range of about 1 weight percent to about 10 weight percent and a concentration of aluminum at the surface of the silicon modified diffusion aluminide layer in the range of about 5 weight percent to about 25 weight percent.

Abstract: A method for making a gas turbine engine turbine blade comprising an airfoil section a platform section, an under platform section, and a dovetail section, the exterior surface of the dovetail section comprising a shank exterior surface and a serrated exterior surface. The blade further includes a silicon-modified diffusion aluminide layer a surface of a turbine blade section selected from the group consisting of the exterior surface of the under platform section, the exterior surface of the dovetail section, and combinations thereof, the aluminide layer having a concentration of silicon at a surface of the aluminide layer in the range of about 1 weight percent to about 10 weight percent and a concentration of aluminum at the surface of the aluminide layer in the range of about 5 weight percent to about 25 weight percent.

Abstract: A modified platinum group metal coating composition comprising a phase having a solid solution face-centered cubic (fcc) crystal structure, rich in platinum group materials. In order to be effective, the platinum group metal coating material was modified based on the chemical composition and chemical activity of the substrate material. The platinum group metal coating material was modified to include, in solid solution, elements of the superalloy substrate, specifically nickel (Ni) and cobalt (Co). Depending on the substrate material, the modified platinum group metal coating material may not even include Ni or Co, but may be modified to include amounts of different elements that are consistent with the chemical composition of the substrate. The modified platinum metal coating material also includes aluminum (Al). The composition may include small amounts of a second phase isolated within the fcc phase matrix.

Abstract: The present invention is a gas turbine engine turbine blade comprising an airfoil section having at least an exterior surface, a platform section having an exterior surface, an under platform section having an exterior surface, and a dovetail section having an exterior surface, the exterior surface of the dovetail section comprising a shank exterior surface and a serrated exterior surface.

Abstract: A coating applied as a two layer system. The outer layer is an oxide of a group IV metal selected from the group consisting of zirconium oxide, hafnium oxide and combinations thereof, which are doped with an effective amount of a lanthanum series oxide. These metal oxides doped with a lanthanum series addition comprises a high weight percentage of the outer coating. As used herein, lanthanum series means an element selected from the group consisting of lanthanum (La), cerium (Ce), praseodymium (Pr), neodymium (Nd), promethium (Pm), samarium (Sm), europium (Eu), gadolinium (Gd), terbium (Tb), dysprosium (Dy), holmium (Ho), erbium (Er), thulium (Tm), ytterbium (Yb), lutetium (Lu) and combinations thereof, and lanthanum series oxides are oxides of these elements. When the zirconium oxide is doped with an effective amount of a lanthanum series oxide, a dense reaction layer is formed at the interface of the outer layer of TBC and the CMAS. This dense reaction layer prevents CMAS infiltration below it.

Abstract: A method for applying a highly porous alumina material that is useful in the hot section of a jet aircraft engine. In order to apply the porous alumina, an aluminum-based metal/alumina material known in the art is first placed onto an aircraft engine component substrate. The aluminum-based metal is then dissolved using a solution that will not affect the alumina or the underlying substrate. The alumina is then washed with deionized water and dried. The aircraft engine component may be first masked by applying a non-porous metal oxide material to the component or by oxidizing the surface of the component. The resulting alumina has a porosity in the range of about 20% to about 45%. The alumina has globular interconnected surface features in the range of about 0.5 ?m to about 20 ?m.

Abstract: A method for applying a highly porous alumina material that is useful in the hot section of a jet aircraft engine. In order to apply the porous alumina, an aluminum-based metal/alumina material known in the art is first placed onto an aircraft engine component substrate. The aluminum-based metal is then dissolved using a solution that will not affect the alumina or the underlying substrate. The alumina is then washed with deionized water and dried. The aircraft engine component may be first masked by applying a non-porous metal oxide material to the component or by oxidizing the surface of the component. The resulting alumina has a porosity in the range of about 20% to about 45%. The alumina has globular interconnected surface features in the range of about 0.5 &mgr;m to about 20 &mgr;m.

Abstract: A highly porous alumina material and a method for applying such material that is useful in the hot section of a jet aircraft engine. In order to manufacture the porous alumina, an aluminum-based metal/alumina material known in the art is first placed onto a metal alloy substrate. The aluminum-based metal is then dissolved using a solution that will not affect the alumina or the underlying substrate. The alumina is then washed with deionized water and dried. The resulting alumina has a porosity in the range of about 20% to about 45%. The alumina has globular interconnected surface features in the range of about 0.5 &mgr;m to about 20 &mgr;m.

Abstract: A starting workpiece of a titanium-base alloy having a temperature-composition phase diagram with a beta region and an alpha-beta region separated by a beta transus temperature is processed by first forging the starting workpiece at a first temperature in the beta region to form a billet, thereafter second forging the billet at a second temperature in the alpha-beta region, thereafter third forging the billet at a third temperature in the beta region, thereafter fourth forging the billet at a fourth temperature in the alpha-beta region so that the step of fourth forging accomplishes a reduction in cross-sectional area of from about 5 to about 40 percent, and thereafter ultrasonic testing the billet. The beta-region third forging step combined with a relatively small reduction during the alpha-beta-region fourth forging step produce a microstructure that is conducive to ultrasonic inspection with minimal interference from noise.

Abstract: An alpha-beta titanium-base alloy is heat treated to improve its dwell fatigue properties while retaining a good balance of mechanical properties. The heat treatment includes first heating the alpha-beta titanium-base alloy to a first heat-treatment temperature in a first range of from about 70° F. below a beta transus temperature of the alpha-beta titanium-base alloy to the beta transus temperature of the alpha-beta titanium-base alloy, and quenching the alpha-beta titanium-base alloy at a rate of greater than about 200° F. per minute. The alpha-beta titanium-base alloy is second heated to a second heat-treatment temperature in a second range of from about 100° F. to about 400° F. below the beta transus temperature of the alpha-beta titanium-base alloy, and thereafter cooling the alpha-beta titanium-base alloy to ambient temperature at a rate of from about 10° F. per minute to about 200° F. per minute.

Abstract: An alpha-beta titanium alloy preform is processed in the beta phase field, by heat treating or beta forging. The processed preform is thereafter heated into the alpha-beta phase field, and a preselected portion is forged, leaving a nonselected portion that is not forged in the alpha-beta phase field. The resulting article has a beta-processed structure in the nonselected portion, and a beta-processed plus alpha-beta forged structure in the preselected portion. In one application, the preform has the shape of a disk useful in the manufacture of an aircraft gas turbine engine. Depending upon specific requirements, either the center or the rim of the disk may be the selected portion.

Abstract: An alpha-beta titanium alloy preform is processed in the beta phase field, by heat treating or beta forging. The processed preform is thereafter heated into the alpha-beta phase field, and a preselected portion is forged, leaving a nonselected portion that is not forged in the alpha-beta phase field. The resulting article has a beta-processed structure in the nonselected portion, and a beta-processed plus alpha-beta forged structure in the preselected portion. In one application, the preform has the shape of a disk useful in the manufacture of an aircraft gas turbine engine. Depending upon specific requirements, either the center or the rim of the disk may be the selected portion.

Abstract: A ferritic alloy steel having good formability, cyclic oxidation resistance and creep strength at elevated temperatures above 1000.degree. F. and particularly above about 1500.degree. F. (816.degree. C.) after a final anneal at 1850.degree. to 2050.degree. F. (1010.degree. to 1120.degree. C.), comprising 0.05% maximum carbon, about 2% maximum manganese, greater than 1.0% to 2.25% silicon, less than 0.5% aluminum, with silicon being at least 3 times the aluminum content, about 6% to about 25% chromium, up to about 5% molybdenum, with the sum of chromium and molybdenum being at least 8%, 0.05% maximum nitrogen, at least one of titanium, zirconium and tantalum, with said titanium, zirconium and tantalum being present in an amount at least equal to the stoichiometric equivalent of the present carbon plus the percent nitrogen, at least 0.1% uncombined columbium, and balance essentially iron.