Abstract: Delamination of thermal barrier coatings (“TBC's”) (276) from superalloy substrates (262) of components (260) for turbine engines (80), such as engine blades (92), vanes (104, 106), or castings in transitions (85), is inhibited during subsequent cooling passage (270) formation. Partially completed cooling passages (264), which have skewed passage paths that end at a terminus (268), which is laterally offset from the passage entrance (266), are formed in the superalloy component (260) prior to application of the TBC layer(s) (276). The skewed, laterally offset path of each partially completed cooling passage (264) establishes an overhanging shield layer (269) of superalloy material that protects the TBC layer (276) during completion of the cooling passage (270).

Abstract: Engineered groove features (EGFs) are formed within thermal barrier coatings (TBCs) of turbine engine components. The EGFs are advantageously aligned with likely stress zones within the TBC or randomly aligned in a convenient two-dimensional or polygonal planform pattern on the TBC surface and into the TBC layer. The EGFs localize thermal stress- or foreign object damage (FOD)-induced crack propagation within the TBC that might otherwise allow excessive TBC spallation and subsequent thermal exposure damage to the turbine component underlying substrate. Propagation of a crack is arrested when it reaches an EGF, so that it does not cross over the groove to otherwise undamaged zones of the TBC layer. In some embodiments, the EGFs are combined with engineered surface features (ESFs) that are formed in the component substrate or within intermediate layers applied between the substrate and the TBC.

Abstract: Delamination of thermal barrier coatings (“TBC's”) (276) from superalloy substrates (262) of components (260) for turbine engines (80), such as engine blades (92), vanes (104, 106), or castings in transitions (85), is inhibited during subsequent cooling passage (270) formation. Partially completed cooling passages (264), which have skewed passage paths that end at a terminus (268), which is laterally offset from the passage entrance (266), are formed in the superalloy component (260) prior to application of the TBC layer(s) (276). The skewed, laterally offset path of each partially completed cooling passage (264) establishes an overhanging shield layer (269) of superalloy material that protects the TBC layer (276) during completion of the cooling passage (270).

Abstract: Cooling passages (99, 105) are formed in components for combustion turbine engines, such as blades (92), vanes (104, 106), ring segments (110) or castings in transitions (85), during investment casting, through use of ceramic shell inserts (130) within the casting mold (152). Ceramic posts (134) formed in the ceramic shell insert (130) have profiles conforming to corresponding profiles of partially completed cooling passages (156). Posts (134) are removed after superalloy component casting, forming the partially completed cooling passages, which are subsequently completed by removing remaining superalloy material along the cooling passage path.

Abstract: A metallic coating or alloy is provided. The metallic coating or alloly includes iron, chromium, aluminum, tantalum, and nickel and contains no rhenium. The presence of tantalum and iron and the absence of rhenium are effective to increase a ?/?? transition temperature of the alloy. A component including the metallic coating or alloy is also provided.

Abstract: A turbine abradable component includes a support surface and a thermally sprayed ceramic/metallic abradable substrate coupled to the support surface for orientation proximal a rotating turbine blade tip circumferential swept path. An elongated pixelated major planform pattern (PMPP) of a plurality of discontinuous micro surface features (MSF) project from the substrate surface. The PMPP repeats radially along the swept path in the blade tip rotational direction, for selectively directing airflow between the blade tip and the substrate surface. Each MSF is defined by a pair of first opposed lateral walls defining a width, length and height that occupy a volume envelope of 1-12 cubic millimeters. The PMPP arrays of MSFs provide airflow control of hot gasses in the gap between the abradable surface and the blade tip with smaller potential rubbing surface area than solid projecting ribs with similar planform profiles.

Abstract: Cooling holes in a turbine component, such as a blade, vane or combustor transition, are formed in and surrounded by a micro surface feature (MSF) that protects the adjoining thermal barrier coating (TBC) from delamination or crack propagation during the hole formation or during engine operation. The MSF effectively functions as a circumferential sleeve around the cooling hole margin so that relatively more friable TBC material that would otherwise define the cooling hole margin is not directly exposed to coolant fluid exhausting the hole, foreign object damage (FOD) or contact with cooling hole formation tooling when fabricating the hole through the TBC layer. The MSF is formed as a projection from the component substrate or during subsequent application of a metallic bond coat (BC) layer.

Abstract: Engineered groove features (EGFs) are formed within thermal barrier coatings (TBCs) of turbine engine components. The EGFs are advantageously aligned with likely stress zones within the TBC or randomly aligned in a convenient two-dimensional or polygonal planform pattern on the TBC surface and into the TBC layer. The EGFs localize thermal stress- or foreign object damage (FOD)-induced crack propagation within the TBC that might otherwise allow excessive TBC spallation and subsequent thermal exposure damage to the turbine component underlying substrate. Propagation of a crack is arrested when it reaches an EGF, so that it does not cross over the groove to otherwise undamaged zones of the TBC layer. In some embodiments, the EGFs are combined with engineered surface features (ESFs) that are formed in the component substrate or within intermediate layers applied between the substrate and the TBC.

Abstract: A thermal barrier tile (34) with a braze layer (46) co-sintered to a ceramic layer (48), optionally with a layer of MCrAlY bond material (74) disposed there between. The tile can be brazed to a substrate (26) of a component for fabrication or repair of a thermal barrier coating (28). The tile may be fabricated by disposing a first layer of a metal brazing material in a die case (40); disposing a second layer of a ceramic powder on the metal brazing material; and co-sintering the two layers with spark plasma sintering to form the co-sintered ceramic/metal tile. The tile may include an interlocking structural pattern (56) at the ceramic/braze interface, and further may include mirror image contoured edges (70) for interlocking of tiles that are disposed side-by-side. Heights of adjacent tiles (34F, 34G) may be different to improve abradability of the surface.

Abstract: A metallic coating or alloy is provided. The metallic coating or alloly includes iron, chromium, aluminum, tantalum, and nickel and contains no rhenium. The presence of tantalum and iron and the absence of rhenium are effective to increase a ?/?? transition temperature of the alloy. A component including the metallic coating or alloy is also provided.

Abstract: A metallic coating is provided. The nickel based metallic coating includes tantalum, cobalt, chromium, and aluminum. The nickel based metallic coating does not include silicon and/or hafnium and/or zirconium. A tantalum addition in nickel based coating stabilized the phases gamma/gamma1 at high temperatures leading to a reduction of local stresses.

Abstract: A method for casting an object (12) having an integrated surface feature (10) for location, inspection, and analysis using a feature-based vision system is provided herein that includes determining a shape geometry for a surface feature (10), wherein the shape geometry is adapted for tracking with a feature-based vision system, determining a proper size, placement, and orientation for the surface feature (10) based on a type of inspection, and casting the surface feature (10) into an object (12) at the determined placement and orientation using an investment casting process to produce an integrated surface feature.

Abstract: A thermal barrier tile (34) with a braze layer (46) co-sintered to a ceramic layer (48) is brazed to a substrate (26) of a component for fabrication or repair of a thermal barrier coating (28) for example on a gas turbine ring segment (22, 24). The tile may be fabricated by disposing a first layer of a metal brazing material in a die case (40); disposing a second layer of a ceramic powder on the metal brazing material; and co-sintering the two layers with spark plasma sintering to form the co-sintered ceramic/metal tile. A material property of an existing thermal barrier coating to be repaired may be determined (90), and the co-sintering may be controlled (93) responsive to the property to produce tiles compatible with the existing thermal barrier coating in a material property such as thermal conductivity.

Abstract: A metallic coating or alloy is provided. The metallic coating includes iron, cobalt, chromium, and aluminum. Tantalum may also be included. A new addition in nickel based coating with stabilized gamma/gamma? phases at high temperatures lead to a reduction of local stresses. A component including the metallic coating or alloy is also provided.

Abstract: A method for forming a textured bond coat surface (48) for a thermal barrier coating system (44) of a gas turbine component (34). The method includes selectively melting portions of a layer of alloy particles (16) with a patterned energy beam (20) to form successive layers of alloy material (16?, 16?) until a desired surface geometric feature (26) is achieved. The energy beam pattern may be indexed between layers to form a protruding undercut (28) in the geometric feature. The patterned energy beam may be formed by directing laser energy from a diode laser (30) through a cartridge filter (32). Particles of a flux material (18) may be melted along with the alloy particles to form a protective layer of slag (22) over the melted and cooling alloy material.

Abstract: Turbine and compressor casing abradable component embodiments for turbine engines, with composite grooves and vertically projecting rows of ridges in planform patterns, establishing upper and lower wear zones. The lower wear zone reduces, redirects and/or blocks blade tip downstream airflow leakage, while the upper wear zone is optimized to minimize blade tip gap and wear by being more easily abradable than the lower zone. An elongated first ridge in the lower wear zone terminates in a continuous surface plateau. A plurality of second ridges or nibs, separated by grooves, project from the plateau, forming the upper wear zone. Each of the second ridges has a planform cross section smaller than the plateau planform cross section and a height smaller than the first ridge height. Some embodiments of the second ridges have spacing, planform cross sections, heights and separating groove dimensions selected for shearing when contacted by turbine blade tips.

Abstract: A thermal barrier tile (34) with a braze layer (46) co-sintered to a ceramic layer (48) is brazed to a substrate (26) of a component for fabrication or repair of a thermal barrier coating (28) for example on a gas turbine ring segment (22, 24). The tile may be fabricated by disposing a first layer of a metal brazing material in a die case (40); disposing a second layer of a ceramic powder on the metal brazing material; and co-sintering the two layers with spark plasma sintering to form the co-sintered ceramic/metal tile. A material property of an existing thermal barrier coating to be repaired may be determined (90), and the co-sintering may be controlled (93) responsive to the property to produce tiles compatible with the existing thermal barrier coating in a material property such as thermal conductivity.

Abstract: A metallic bondcoat with phases of ? and ?? is provided. The metallic coating or alloy is nickel based. The metallic coating or alloy has ? and ?? phases and optionally has ?-phase. The new addition in nickel based coating stabilizes the phases ? and ?? at high temperatures leading to a reduction of local stresses.

Abstract: A metallic coating or alloy is provided, which is nickel based, and includes at least ? and ?? phases. The metallic coating or the alloy further includes tantalum (Ta) in the range of between 4 wt % to 7.5 wt %. The metallic coating or the alloy also includes cobalt (Co) in the range between 11 wt %-14.5 wt %.

Abstract: A metallic coating is provided. The nickel based metallic coating includes tantalum, cobalt, chromium, and aluminum. The nickel based metallic coating does not include silicon and/or hafnium and/or zirconium. A tantalum addition in nickel based coating stabilized the phases gamma/gamma1 at high temperatures leading to a reduction of local stresses.