Abstract: A method of forming an interstage seal including removing a diaphragm seal box (14) from a gas turbine compressor assembly (10) and removing a labyrinth sealing member (12) from the diaphragm seal box (14). An abradable material layer (34) may be deposited on the diaphragm seal box (14). A spray gun may be mounted in relation to an engine disk (16) of the gas turbine compressor assembly (10) for cold-spraying a quantity of particles toward the engine disk (16). The particles may be sprayed at a velocity sufficiently high to cause at least a portion of the quantity of particles to adhere to the engine disk (16). The spray gun may be controlled to deposit a quantity of particles on the compressor disk (16) to form a geometry (32) that will abrade the abradable material layer (34) during operation of the gas turbine compressor assembly (10). The geometry (32) abrading the abradable material layer (34) forms an interstage seal.

Abstract: One embodiment of a transition-to-turbine seal (300) comprises a first, flattened section (302) adapted to be received in a peripheral axial slot (320) of a transition (325), and a second, generally C-shaped section (301). The generally C-shaped section (301) comprises a flattened portion (305) near the first, flattened section (302), and a curved portion (306) extending to a free edge (307). A fiber metal strip component (309) may be attached to the flattened portion (305) to define a first engagement surface adapted to engage an upstream side (336) of an outer vane seal rail (337), and a second engagement surface 308, adjacent the free edge (307), provides an opposed wear surface adapted to engage a downstream side (338) of the outer vane seal rail (337). System embodiments also are described, in which such transition-to-turbine seal (300) is isolated from a hot gas path (350) by provision of a plurality of cooling apertures (327) in the transition (325).

Abstract: This invention relates to an abradable coating system for use in axial turbine engines. When coated onto a turbine ring seal segment the coating system may allow formation of an individualized seal between turbine blade disks and the surrounding ring seal without causing excessive wear to the blade tips. The abradable coating system includes columns of an abradable material. Thus, interference between the blades and the abradable coating system causes the individual columns to break off at the base. This abrasion mechanism may reduce blade wear and spalling of the coating system when compared to conventional coatings.

Abstract: A combustor device and transition duct assembly is provided for use in a gas turbine engine. The combustor device comprises combustor structure having an exit portion; spring clips mounted to the exit portion of the combustor structure; and a burner assembly. The transition duct comprises a conduit having inlet and outlet sections and an abradable material layer provided along a circumferential portion of the inlet section of the transition duct conduit. The transition duct conduit inlet section may be coupled to the combustor structure exit portion such that the spring clips engage the abradable material layer.

Abstract: A process (20) of producing an abradable thermal barrier coating (38) with a uniformly distributed solid lubricant (46) such as for gas turbine shroud ring segments. A mixture of a ceramic precursor powder (23) and a solid lubricant precursor powder in a liquid (22) is injected (28) into a thermal jet (32), such as one generated by a plasma gun (30). The precursor powders (23, 24) are dissolved or suspended in the liquid, forming a uniform spray mix (34) that is atomized in the thermal jet (32). The resulting spray mix (34) is directed toward a substrate (36), producing layers (39a-39c) of overlapping adherent ceramic splats (42) surrounded generally uniformly by an amount of solid lubricant (46) sufficient to increase abradability of the layers to a given degree, and thus protect turbine blade tips from damage during adjacent rotation with zero clearance.

Abstract: Aspects of the invention are directed to a system for improving engine performance by reducing vane tip clearances in the compressor section of a turbine engine. According to aspects of the invention, an abrasive material can be attached to a rotor disc. Thus, if a vane tip contacts the rotor disc during engine operation, the vane tip is worn away by the abrasive material. The system can reduce the risk of substantial component damage resulting from a vane tip rubbing event. The abrasive material can be provided in the form of an insert that can be removably attached to the rotor disc. Thus, the insert can be readily removed and replaced when necessary.

Abstract: A layered abradable thermal barrier coating (TBC) 20 that is stable, abradabie, and sinter resistant up to about 1400° C. A tungsten bronze structured ceramic of the form Ba6?3xRE8+2xTi18O54, where 0<x<1.5, and RE represents a rare earth lanthanide cation, is applied as a topcoat over a yttria stabilized zirconia (YSZ) undercoat (18) on a bond-coated (17) superalloy metal structure (16). The tungsten bronze structure provides abradability and thermal conductivity. The YSZ layer is a proven concept for thermal barrier coatings, and has demonstrated better adhesion than newer chemistries This combination of layers has synergy that takes advantage of both materials to provide an abradable coating with an extended lifespan on a superalloy substrate compared to prior coatings. The topcoat may be applied with fugitive inclusions that produce porosity to increase abradabilty for improved blade tip clearance control in the turbine section of gas turbines.

Abstract: This invention relates to an abradable coating system for use in axial turbine engines. When coated onto a turbine ring seal segment the coating system may allow formation of an individualized seal between turbine blade disks and the surrounding ring seal without causing excessive wear to the blade tips. The abradable coating system includes columns of an abradable material. Thus, interference between the blades and the abradable coating system causes the individual columns to break off at the base. This abrasion mechanism may reduce blade wear and spalling of the coating system when compared to conventional coatings.

Abstract: The invention relates to a coating system (2) for a component (1) which comprises a porous layer (3) and an abradable layer (4) on the porous layer (3). Further the invention relates to an assembly of two components (1) which are relatively movable to each other and form a gap in between. One component (1) is provided with a coating system (2) and the other component (1) is in sliding contact with the coating system (2).

Abstract: A conformal seal (20) for sealing air flow between a cooling airflow path and a hot gas flow path within a combustion turbine engine. The conformal seal (20) may be fitted within cooperating side slots of adjacent vane segments (10) within the combustion turbine engine. The conformal seal (20) may include an elongated metallic substrate (22, 40) forming an upper surface and a lower surface. A conformal coating (26, 44) may be deposited over one or both surfaces of the substrate (22, 40). The conformal coating (26, 44) may be deposited to a depth so that a point contact between the conformal coating (26, 44) and respective interior walls of the side slots wears the conformal coating (26, 44) to establish surface area contact there between. The surface area contact improves a sealing function between the conformal coating (26, 44) and the respective interior walls during operation of the combustion turbine engine.

Abstract: A bimetallic bond layer (26, 28) for a thermal barrier coating or TBC (30) on a superalloy substrate (22) for a high temperature environment. An interlayer (26) is applied on the substrate. A bond coat (28) comprising a CoNiCrAlY or NiCoCrAlY alloy is applied on the interlayer. A ceramic TBC (30) such as 8YSZ is applied on the bond coat. The interlayer (26) is an alloy that is compatible with the substrate and the bond coat, and that blocks or delays diffusion of aluminum from the bond coat into the substrate at high operating temperatures. This preserves aluminum in the bond coat that maintains a beneficial alumina scale (29) between the bond coat and the TBC. This delays spalling of the TBC, and lengthens the coating and component life.

Abstract: A method for welding to a superalloy material without causing cracking of the material. The method includes the steps of depositing a weld strip (18) of a weldable material onto a superalloy substrate material (12) using a spray deposition process (20) and then forming a weldment (26) to the weld strip. None or a controlled amount of the substrate material (12) is melted during the weld in order to maintain the concentration of strengthening elements in the local melt within a zone of weldability. The spray deposition process may be a thermal process such as HVOF or a cold spray process. A groove (16) may be formed in a surface (10) of the superalloy substrate material to receive the weld strip. A diffusion heat treatment step may be used to improve adhesion between the weld strip and the superalloy material.

Abstract: One embodiment of a transition-to-turbine seal (300) comprises a first, flattened section (302) adapted to be received in a peripheral axial slot (320) of a transition (325), and a second, generally C-shaped section (301). The generally C-shaped section (301) comprises a flattened portion (305) near the first, flattened section (302), and a curved portion (306) extending to a free edge (307). A fiber metal strip component (309) may be attached to the flattened portion (305) to define a first engagement surface adapted to engage an upstream side (336) of an outer vane seal rail (337), and a second engagement surface 308, adjacent the free edge (307), provides an opposed wear surface adapted to engage a downstream side (338) of the outer vane seal rail (337). System embodiments also are described, in which such transition-to-turbine seal (300) is isolated from a hot gas path (350) by provision of a plurality of cooling apertures (327) in the transition (325).

Abstract: A conformal seal (20) for sealing air flow between a cooling airflow path and a hot gas flow path within a combustion turbine engine. The conformal seal (20) may be fitted within cooperating side slots of adjacent vane segments (10) within the combustion turbine engine. The conformal seal (20) may include an elongated metallic substrate (22, 40) forming an upper surface and a lower surface. A conformal coating (26, 44) may be deposited over one or both surfaces of the substrate (22, 40). The conformal coating (26, 44) may be deposited to a depth so that a point contact between the conformal coating (26, 44) and respective interior walls of the side slots wears the conformal coating (26, 44) to establish surface area contact there between. The surface area contact improves a sealing function between the conformal coating (26, 44) and the respective interior walls during operation of the combustion turbine engine.

Abstract: Aspects of the invention are directed to a system for improving engine performance by reducing vane tip clearances in the compressor section of a turbine engine. According to aspects of the invention, an abrasive material can be attached to a rotor disc. Thus, if a vane tip contacts the rotor disc during engine operation, the vane tip is worn away by the abrasive material. The system can reduce the risk of substantial component damage resulting from a vane tip rubbing event. The abrasive material can be provided in the form of an insert that can be removably attached to the rotor disc. Thus, the insert can be readily removed and replaced when necessary.

Abstract: An abradable thermal barrier coating material (10) formed of a highly defective fluorite ceramic matrix (16) having a desired degree of porosity (18) created in part by the addition of a fugitive material (19). The ceramic material has a concentration of a stabilizer sufficiently high that the oxygen vacancies created by the stabilizer interact within the matrix to form multi-vacancies, thereby improving the sintering resistance of the material. Such a concentration of stabilizer results in a material that is softer than prior art materials having lower concentrations of stabilizer, and that will be more resistive to sintering than prior art materials. Embodiments include a fluorite matrix of zirconia stabilized by at least 30 wt. % yttria, or stabilized by at least 30 wt. % ytterbia, and with porosity of 10-40%. In one embodiment, a metallic gas turbine seal ring segment is coated with a bond coat layer, then with a layer of porous 8 wt. % YSZ material, and finally with a layer of 33 mole % YbSZ (61.3 wt.

Abstract: A device operable in a temperature environment in excess of about 1000° C. is provided. The device comprises a substrate and a ceramic thermal barrier layer deposited on at least a portion of the substrate. The layer is formed with a ternary or pseudoternary oxide having a pyrochlore or perovskite structure and a fugative material and having pores or other voluminous defects. The thermal barrier layer advantageously is abradable.

Abstract: A fusion weldable superalloy containing 0.005-0.5 wt. % scandium. In one embodiment, the superalloy may have a composition similar to IN-939 alloy, but having added scandium and having only 0.005-0.040 wt. % zirconium. A gas turbine component may be formed by an investment casting of such a scandium-containing superalloy, and may include a fusion weld repaired area. A scandium-containing nickel-based superalloy coated with an MCrAlY bond coat will have improved cyclic oxidation resistance due to the sulfur-gettering effect of the scandium.

Abstract: An abradable thermal barrier coating material (10) formed of a highly defective fluorite ceramic matrix (16) having a desired degree of porosity (18) created in part by the addition of a fugitive material (19). The ceramic material has a concentration of a stabilizer sufficiently high that the oxygen vacancies created by the stabilizer interact within the matrix to form multi-vacancies, thereby improving the sintering resistance of the material. Such a concentration of stabilizer results in a material that is softer than prior art materials having lower concentrations of stabilizer, and that will be more resistive to sintering than prior art materials. Embodiments include a fluorite matrix of zirconia stabilized by at least 30 wt. % yttria, or stabilized by at least 30 wt. % ytterbia, and with porosity of 10-40%. In one embodiment, a metallic gas turbine seal ring segment is coated with a bond coat layer, then with a layer of porous 8 wt. % YSZ material, and finally with a layer of 33 mole % YbSZ (61.3 wt.

Abstract: A fusion weldable superalloy containing 0.005-0.5 wt. % scandium. In one embodiment, the superalloy may have a composition similar to IN-939 alloy, but having added scandium and having only 0.005-0.040 wt. % zirconium. A gas turbine component may be formed by an investment casting of such a scandium-containing superalloy, and may include a fusion weld repaired area. A scandium-containing nickel-based superalloy coated with an MCrAlY bond coat will have improved cyclic oxidation resistance due to the sulfur-gettering effect of the scandium.