Abstract: Thermal barrier coatings consist of a tantala-zirconia mixture that is stabilized with two or more stabilizers. An exemplary thermal barrier coating consists of, by mole percent: about 8% to about 30% YO1.5; about 8% to about 30% YbO1.5 or GdO1.5 or combination thereof; about 8% to about 30% TaO2.5; about 0% to about 10% HfO2; and a balance of ZrO2.

Abstract: Systems and methods are provided for cleaning one or more cooling passages associated with a combustion chamber of a gas turbine engine. The gas turbine engine has a compressor section upstream from the combustion section. The method includes receiving a pressurized fluid from a source and directing the pressurized fluid through an inlet of a chamber such that a portion of a plurality of particles within the chamber is entrained within the pressurized fluid. The method includes injecting the pressurized fluid with the entrained portion of the plurality of particles downstream from the compressor section through an end wall of a diffuser and deswirl system upstream from a combustor plenum of the combustion section to clean the one or more cooling passages associated with the combustion chamber.

Abstract: A method of forming an abrasive nickel-based alloy on a turbine blade tip includes producing or obtaining a metal powder that is mixed with a carbon powder to form a carbon-enriched metal powder. The metal powder includes a refractory element. The method further includes bonding the carbon-enriched metal powder to the turbine blade tip. The step of bonding includes raising the temperature of the carbon-enriched metal powder past its melting point, thereby causing the carbon to combine with the refractory elements to form abrasive carbide particles.

Abstract: Thermal barrier coatings, which may be used in gas turbine engines, comprise or consist of a tantala-niobia-zirconia mixture that is stabilized with two or more stabilizers. An exemplary thermal barrier coating comprises or consists of, by mole percent: about 2% to about 30% YO1.5; about 8% to about 30% YbO1.5 or GdO1.5 or combination thereof; about 6% to about 30% TaO2.5; about 0.1% to about 10% NbO2.5; about 0% to about 10% HfO2; and a balance of ZrO2.

Abstract: Systems and methods are provided for cleaning one or more cooling passages associated with a combustion chamber of a gas turbine engine. The gas turbine engine has a compressor section upstream from the combustion section. The method includes receiving a pressurized fluid from a source and directing the pressurized fluid through an inlet of a chamber such that a portion of a plurality of particles within the chamber is entrained within the pressurized fluid. The method includes injecting the pressurized fluid with the entrained portion of the plurality of particles downstream from the compressor section through an end wall of a diffuser and deswirl system upstream from a combustor plenum of the combustion section to clean the one or more cooling passages associated with the combustion chamber.

Abstract: A method of forming an abrasive nickel-based alloy on a turbine blade tip includes producing or obtaining a metal powder that is mixed with a carbon powder to form a carbon-enriched metal powder. The metal powder includes a refractory element. The method further includes bonding the carbon-enriched metal powder to the turbine blade tip. The step of bonding includes raising the temperature of the carbon-enriched metal powder past its melting point, thereby causing the carbon to combine with the refractory elements to form abrasive carbide particles.

Abstract: Systems and methods are provided for cleaning one or more cooling passages associated with a combustion chamber of a gas turbine engine. The gas turbine engine has a compressor section upstream from the combustion chamber. In one embodiment, a method includes: receiving a pressurized fluid from a source; directing the pressurized fluid through an inlet of a chamber such that a portion of a plurality of particles within the chamber is entrained within the pressurized fluid; and injecting the pressurized fluid with the entrained portion of the plurality of particles downstream from the compressor section through an end wall of a diffuser and deswirl system of the gas turbine engine or through a plenum upstream from the combustion chamber to clean the one or more cooling passages associated with the combustion chamber.

Abstract: Thermal barrier coatings, which may be used in gas turbine engines, comprise or consist of a tantala-niobia-zirconia mixture that is stabilized with two or more stabilizers. An exemplary thermal barrier coating comprises or consists of, by mole percent: about 2% to about 30% YO1.5; about 8% to about 30% YbO1.5 or GdO1.5 or combination thereof; about 6% to about 30% TaO2.5; about 0.1% to about 10% NbO2.5; about 0% to about 10% HfO2; and a balance of ZrO2.

Abstract: Systems and methods are provided for cleaning one or more cooling passages associated with a combustion chamber of a gas turbine engine. The gas turbine engine has a compressor section upstream from the combustion chamber. In one embodiment, a method includes: receiving a pressurized fluid from a source; directing the pressurized fluid through an inlet of a chamber such that a portion of a plurality of particles within the chamber is entrained within the pressurized fluid; and injecting the pressurized fluid with the entrained portion of the plurality of particles downstream from the compressor section through an end wall of a diffuser and deswirl system of the gas turbine engine or through a plenum upstream from the combustion chamber to clean the one or more cooling passages associated with the combustion chamber.

Abstract: A turbine engine, an engine structure, and a method of forming an engine structure are provided herein. In an embodiment, an engine structure includes a metal substrate, a thermal barrier coating layer, and a metal silicate protective layer. The thermal barrier coating layer overlies the metal substrate, and the thermal barrier coating layer has columnar grains with gaps defined between the columnar grains. The metal silicate protective layer is formed over the thermal barrier coating layer, and the metal silicate protective layer covers the columnar grains and the gaps between the columnar grains.

Abstract: Methods of forming dispersoid hardened metallic materials are provided. In an exemplary embodiment, a method of producing dispersoid hardened metallic materials includes forming a starting composition with a base metal component and a dispersoid forming component. The starting composition includes the base metal component in an amount from about 50 to about 99.999 weight percent and the dispersoid forming component in an amount from about 0.001 to about 1 weight percent, based on the total weight of the starting composition. A starting powder is formed from the starting composition, and the starting powder is fluidized with a fluidizing gas for a period of time sufficient to oxidize the dispersoid forming component to form the dispersoid hardened metallic material. The dispersoid forming component is oxidized while the starting powder is a solid.

Abstract: A ceramic composition of matter includes a solvent material and a solute material. The solvent material comprises zirconia or hafnia, or a mixture thereof, which is stabilized with a stabilizing oxide. The solute material comprises a metal oxide of the formula, XaOb, where X is a metallic element, O is oxygen, and “a” and “b” are stoichiometric coefficients. The difference in density between the solvent material and the solute material is less than about 30%.

Abstract: Methods of forming dispersoid hardened metallic materials are provided. In an exemplary embodiment, a method of producing dispersoid hardened metallic materials includes forming a starting composition with a base metal component and a dispersoid forming component. The starting composition includes the base metal component in an amount from about 50 to about 99.999 weight percent and the dispersoid forming component in an amount from about 0.001 to about 1 weight percent, based on the total weight of the starting composition. A starting powder is formed from the starting composition, and the starting powder is fluidized with a fluidizing gas for a period of time sufficient to oxidize the dispersoid forming component to form the dispersoid hardened metallic material. The dispersoid forming component is oxidized while the starting powder is a solid.

Abstract: Thermal barrier coatings consist of a tantala-zirconia mixture that is stabilized with two or more stabilizers. An exemplary thermal barrier coating consists of, by mole percent: about 8% to about 30% YO1.5; about 8% to about 30% YbO1.5 or GdO1.5 or combination thereof; about 8% to about 30% TaO2.5; about 0% to about 10% HfO2; and a balance of ZrO2.

Abstract: A turbine engine, an engine structure, and a method of forming an engine structure are provided herein. In an embodiment, an engine structure includes a metal substrate, a thermal barrier coating layer, and a metal silicate protective layer. The thermal barrier coating layer overlies the metal substrate, and the thermal barrier coating layer has columnar grains with gaps defined between the columnar grains. The metal silicate protective layer is formed over the thermal barrier coating layer, and the metal silicate protective layer covers the columnar grains and the gaps between the columnar grains.

Abstract: Embodiments of methods and systems for inspecting a structure for a crystallographic imperfection are provided. In the method, an X-ray wavelength that is particularly susceptible to diffraction by the crystallographic imperfection is identified. Then an X-ray source is provided to emit X-rays in the identified X-ray wavelength. While placing the structure at a sequence of positions relative to the X-ray source, X-rays are directed at the structure in multiple, non-parallel arrays to create sequential patterns of diffracted X-rays. The patterns of diffracted X-rays are digitally captured and communicated to a computer that compares them to locate the crystallographic imperfection. For a surface imperfection, the imperfection may be marked with a target to allow for physical removal.

Abstract: A thermal barrier coating is formed over the substrate. A majority of the thermal barrier coating comprises a multi-phase material comprising a polycrystalline material including two or more phases. Each phase forms an individual grain, adjacent individual grains are separated by grain boundaries, each phase comprises an oxide compound, the multi-phase material is formed from three or more constituents, the three or more constituents consist of different materials that are not completely soluble in each other, and the two or more phases are not completely soluble in each other and do not form only one compound.

Abstract: Embodiments of methods and systems for inspecting a structure for a crystallographic imperfection are provided. In the method, an X-ray wavelength that is particularly susceptible to diffraction by the crystallographic imperfection is identified. Then an X-ray source is provided to emit X-rays in the identified X-ray wavelength. While placing the structure at a sequence of positions relative to the X-ray source, X-rays are directed at the structure in multiple, non-parallel arrays to create sequential patterns of diffracted X-rays. The patterns of diffracted X-rays are digitally captured and communicated to a computer that compares them to locate the crystallographic imperfection. For a surface imperfection, the imperfection may be marked with a target to allow for physical removal.

Abstract: Turbine components are provided. In an embodiment, by way of example, a hub and a ring are included. The hub comprises a first material. The ring is bonded to the hub. The ring comprises a plurality of arc segments forming a ring, each arc segment comprising a second material comprising a single crystal superalloy material having a predetermined primary orientation and a predetermined secondary orientation, each predetermined primary orientation of each arc segment being substantially equal, and each predetermined secondary orientation of each arc segment being substantially equal, each arc segment adjacent another arc segment, and the adjacent arc segments having a predetermined crystallographic mismatch therebetween. Methods of manufacturing are also provided.

Abstract: Coated components and methods of fabricating coated components and coated turbine disks are provided. In an embodiment, by way of example only, a coated component includes a substrate comprising a superalloy in an unmodified form and a coating disposed over the substrate, where the coating comprises the superalloy in a modified form.