Abstract: A ceramic insulating composition including monoclinic zirconia-hafnia. The zirconia-hafnia may be used as a filler powder (14) in a ceramic composition (10), or it may be used as a protective overlayer (50). The zirconia-hafnia (24) may be combined with alumina (22) to achieve a desired elastic modulus. The portion of hafnia contained in the zirconia-hafnia is selected to provide a predetermined phase transition temperature to minimize undesired phase changes during high temperature use of the material.

Abstract: A material (100) formed of a sintered aggregation of ceria particles (106), mullite particles (108) and an alumina matrix material (110). Differential thermal expansion of the ceria and mullite particles generates thermal stress sufficient to create micro-cracking of the ceria particles. The ratio of ceria to mullite may be selected to achieve a desired coefficient of thermal expansion for matching the thermal growth of a mating CMC material (102). The micro cracks provide the material with a desired degree of strain tolerance useful in high temperature applications such as a solid core gas turbine vane (20).

Abstract: A xenotime phosphate protective overlayer (22) for protecting a ceramic material (24) from a high temperature, moisture-containing environment. Yttrium phosphate may be used as a protective overlayer to protect an underlying mullite layer to temperatures in excess of 1,500° C. The coating may have porosity of greater than 15% for improved thermal shock protection. To prevent the ingress of oxygen to an underlying ceramic non-oxide material, such as silicon carbide or silicon nitride, an oxygen barrier layer (34) is disposed between the xenotime phosphate coating and the non-oxide material. Such a protective overlayer may be used for an article having a ceramic matrix composite substrate.

Abstract: An article comprising a ceramic matrix composite component of a gas turbine engine such as a combustion liner this is adapted for backside radiative cooling is provided. The article comprises a ceramic matrix composite composition having a frontside and a backside, and coated with a ceramic insulating material on the frontside and coated with a high temperature emissive material on the backside; and a metal element spaced apart from the ceramic matrix composition and defining a gap between the metal element and the ceramic matrix composite, whereby at least a portion of thermal energy exposed to the ceramic insulating material is emitted from the high temperature emissive material through the gap to the metal element.

Abstract: A multi-layer ceramic matrix composite structure (40) having a plurality of fiber-reinforced cooling passages (42) formed therein. The cooling passages are formed by the removal of a fugitive material (24). The fugitive material is part of a wrapped fugitive material structure (28) containing a layer of reinforcing ceramic fibers (26) that is used to lay-up the multi-layer structure. An intermediate layer of ceramic fabric 56 may be placed alternately over and under the wrapped fugitive material structure to separate the cooling passages into alternating upper (54) and lower (52) cooling passages. The transversely oriented fibers surrounding the cooling passages serve to increase the interlaminar strength of the structure when compared to prior art designs. An airfoil member (112) incorporating such reinforced integral cooling passages (120) is provided.

Abstract: A material (100) formed of a sintered aggregation of ceria particles (106), mullite particles (108) and an alumina matrix material (110). Differential thermal expansion of the ceria and mullite particles generates thermal stress sufficient to create micro-cracking of the ceria particles. The ratio of ceria to mullite may be selected to achieve a desired coefficient of thermal expansion for matching the thermal growth of a mating CMC material (102). The micro cracks provide the material with a desired degree of strain tolerance useful in high temperature applications such as a solid core gas turbine vane (20).

Abstract: A combustion catalyst coating (36) applied to the surface of a ceramic thermal barrier coating (34) which is supported by a metal substrate (32). The columnar-grained microstructure of the thermal barrier coating surface provides the necessary surface area for interaction of the catalyst and a fuel-air mixture in a catalytic combustor of a gas turbine engine. The temperature gradient developed across the thermal barrier coating protects the underlying metal substrate from a high temperature combustion process occurring at the catalyst surface. The thermal barrier coating deposition process may be controlled to form a columnar grained microstructure having a plurality of primary columns each with a plurality of secondary and tertiary branches in order to achieve a desired specific surface area for receiving the catalyst coating.

Abstract: A multi-layer ceramic matrix composite structure (40) having a plurality of fiber-reinforced cooling passages (42) formed therein. The cooling passages are formed by the removal of a fugitive material (24). The fugitive material is part of a wrapped fugitive material structure (28) containing a layer of reinforcing ceramic fibers (26) that is used to lay-up the multi-layer structure. An intermediate layer of ceramic fabric 56 may be placed alternately over and under the wrapped fugitive material structure to separate the cooling passages into alternating upper (54) and lower (52) cooling passages. The transversely oriented fibers surrounding the cooling passages serve to increase the interlaminar strength of the structure when compared to prior art designs. An airfoil member (112) incorporating such reinforced integral cooling passages (120) is provided.

Abstract: A composite (20) has a matrix (10), preferably of ceramic, interspersed with reinforcement structures (12), preferably ceramic fibers, coated with a material (14) selected from ZrGeO4, HfGeO4 and CeGeO4, where the composite can be used as a component in high temperature turbines.

Abstract: A xenotime phosphate protective overlayer (22) for protecting a ceramic material (24) from a high temperature, moisture-containing environment. Yttrium phosphate may be used as a protective overlayer to protect an underlying mullite layer to temperatures in excess of 1,500° C. The coating may have porosity of greater than 15% for improved thermal shock protection. To prevent the ingress of oxygen to an underlying ceramic non-oxide material, such as silicon carbide or silicon nitride, an oxygen barrier layer (34) is disposed between the xenotime phosphate coating and the non-oxide material. Such a protective overlayer may be used for an article having a ceramic matrix composite substrate.

Abstract: A vane assembly for a turbine assembly includes an inner endcap, an outer endcap, and a body. The body includes a metallic core assembly, a ceramic shell assembly and a support assembly. The metallic core assembly is coupled to the inner and outer endcaps and bears most of the mechanical loads, including aerodynamic loads. The ceramic shell bears substantially all of the thermal stress placed on the vane assembly. The support assembly is disposed between the metallic core assembly and said ceramic shell assembly and is coupled to the metallic core assembly.