Turbine engine (80) components, such as blades (92), vanes (104, 106), ring segment 110 abradable surfaces 120, or transitions (85), have furcated engineered groove features (EGFs) (403, 404, 418, 509, 511, 512) that cut into the outer surface of the component's thermal barrier coating (TBC). In some embodiments, the EGF planform pattern defines adjoining outer hexagons (560, 640, 670, 690, 710). In some embodiments, the EGF pattern further defines within each outer hexagon (560, 640, 670, 690, 710) a planform pattern of adjoining inner polygons (570, 580, 590, 600, 610, 680, 682, 700, 702, 704, 705, 720). At least three respective groove segments (509, 511, 512) within the EGF pattern (506, 507, 508) converge at each respective outer hexagonal vertex (510, 564) or inner polygonal vertex (574, 564, 604, 614) in a multifurcated pattern, so that crack-inducing stresses are attenuated in cascading fashion, as the stress (σ.sub.A) is furcated (σ.sub.B, σ.sub.C) at each successive vertex juncture.

An article includes a substrate, a ceramic barrier coating, and a layer of networked ceramic nanofibers. The ceramic barrier coating is disposed on the substrate and has a porous columnar microstructure. The layer of networked ceramic nanofibers is disposed on the ceramic barrier layer and seals the pores of the porous columnar microstructure.

A thermal barrier coating disposed on a substrate comprising a plurality of surface features formed on the substrate proximate an inner side of the substrate, each of the plurality of surface features comprising a metallic column having a top with rounded edges; a dense layer disposed in a valley located between each of the plurality of surface features, and the dense layer disposed on the top and covering the rounded edges; a thermally insulating topcoat disposed over the plurality of surface features.

A nozzle including a vane and a band, each having defined therein interlocking features. The vane and the band are each formed of a ceramic matrix composite (CMC) including reinforcing fibers embedded in a matrix. The vane and the band include one or more interlocking features. The nozzle further including an interlocking mechanical joint joining the vane and the band to one another. Methods are also provided for joining the vane and the band at the interlocking features to form an interlocking mechanical joint.

A turbine system including a sealing component is presented. The sealing component includes a ceramic material. The ceramic material includes grains having an average grain size of less than 10 microns. A turbine shroud assembly including the sealing component is also presented.

High temperature stable thermal barrier coatings useful for substrates that form component parts of engines such as a component from a gas turbine engine exposed to high temperatures are provided. The thermal barrier coatings include a multiphase composite and/or a multilayer coating comprised of two or more phases with at least one phase providing a low thermal conductivity and at least one phase providing mechanical and erosion durability. Such low thermal conductivity phase can include a rare earth zirconate and such mechanical durability phase can include a rare earth a rare earth aluminate. The different phases are thermochemically compatible even at high temperatures above about 1200° C.

A method of localized repair to a damaged thermal barrier, the method including subjecting a part coated in a damaged thermal barrier to electrophoresis treatment, the part being made of an electrically conductive material, the damaged thermal barrier including a ceramic material and presenting at least one damaged zone that is to be repaired, the part being present in an electrolyte including a suspension of particles in a liquid medium, the ceramic coating being deposited by electrophoresis in the damaged zone in order to obtain a repaired thermal barrier for use at temperatures higher than or equal to 1000° C., the particles being made of a material different from the ceramic material present in the damaged thermal barrier.

A turbocharger system may include a turbine housing and a tongue insert. The turbine housing may include an inlet, an outlet, and a gas pathway between the inlet and outlet. The gas pathway may include a volute portion and an inlet portion extending approximately tangent to the volute portion. The turbine housing may be formed from a first material. The tongue insert may be received in the turbine housing and may at least partially define the volute portion and the inlet portion. The tongue insert may be formed from a second material that is more heat resistant than the first material.

Provided is a light activated rotor comprising typically a plurality of vanes affixed to a hub rotatable around the longitudinal axis of an axle. Each vane comprises a planar surface oriented generally perpendicular to the longitudinal axis of the axle with each vane separated into a first surface and a second surface. The first and second surface are adjacent and share a common boundary generally perpendicular to the longitudinal axis of the axle. Additionally, the first and second surfaces have differing emissivities. When the light activated rotor is illuminated with a radiant flux, the differing emissivities of the first and second surfaces produce a temperature gradient across the vane and generally perpendicular to the longitudinal axis, and a thermal creep force across the planar surface of the vane generates a revolution of the vane and the affixed hub around the longitudinal axis of the axle.

A turbine component is configured to be cooled by structured porosity cooling. The component includes: a wall; a contiguous porous layer that is part of the wall; a first zone defined in the porous layer such that it has a first structured porosity, and a second zone defined in the layer such that it has a second structured porosity. The first structured porosity is different from the second structured porosity.