Abstract: An airfoil for use in a turbomachine includes a pressure sidewall and a suction sidewall coupled to the pressure sidewall. The suction sidewall and the pressure sidewall define a leading edge and a trailing edge, the leading edge and the trailing edge define a chord distance. The airfoil includes a tip portion extending between the pressure sidewall and the suction sidewall. The tip portion includes a planar section and a recessed section. The recessed section extends adjacent to the planar section such that a thickness of the planar section is less than a thickness of the airfoil. The recessed section is offset a predetermined distance from the leading edge and the trailing edge along the chord distance.

Abstract: Coating systems for components of a gas turbine engine, such as a compressor blade tip, are provided. The coating system can include an abradable material disposed along the compressor blade tip and may be used with a bare compressor casing. The abradable coating is softer than the compressor casing and can reduce the overall rub ratio thereby increasing the lifetime of the compressor blade and casing. Methods are also provided for applying the coating system onto a compressor blade.

Abstract: An airfoil for use in a turbomachine includes a pressure sidewall and a suction sidewall coupled to the pressure sidewall. The suction sidewall and the pressure sidewall define a leading edge and an opposite trailing edge. The leading edge and the trailing edge define a chord distance. The airfoil further includes a root portion, and a tip portion. The tip portion extends between the pressure sidewall and the suction sidewall such that the tip portion is substantially perpendicular to each sidewall. The tip portion includes at least one planar section and at least one oblique section that forms a recess within the tip portion. The at least one oblique section extends from the at least one planar section towards the root portion along the chord distance. The tip portion is configured to reduce airfoil wear during contact with a surrounding casing.

Abstract: Coating systems for components of a gas turbine engine, such as a compressor blade tip, are provided. The coating system can include an abradable material disposed along the compressor blade tip and may be used with a bare compressor casing. The abradable coating is softer than the compressor casing and can reduce the overall rub ratio thereby increasing the lifetime of the compressor blade and casing. Methods are also provided for applying the coating system onto a compressor blade.

Abstract: An airfoil for use in a turbomachine includes a pressure sidewall and a suction sidewall coupled to the pressure sidewall. The suction sidewall and the pressure sidewall define a leading edge and a trailing edge, the leading edge and the trailing edge define a chord distance. The airfoil includes a tip portion extending between the pressure sidewall and the suction sidewall. The tip portion includes a planar section and a recessed section. The recessed section extends adjacent to the planar section such that a thickness of the planar section is less than a thickness of the airfoil. The recessed section is offset a predetermined distance from the leading edge and the trailing edge along the chord distance.

Abstract: An airfoil for use in a turbomachine includes a pressure sidewall and a suction sidewall coupled to the pressure sidewall. The suction sidewall and the pressure sidewall define a leading edge and an opposite trailing edge. The leading edge and the trailing edge define a chord distance. The airfoil further includes a root portion, and a tip portion. The tip portion extends between the pressure sidewall and the suction sidewall such that the tip portion is substantially perpendicular to each sidewall. The tip portion includes at least one planar section and at least one oblique section that forms a recess within the tip portion. The at least one oblique section extends from the at least one planar section towards the root portion along the chord distance. The tip portion is configured to reduce airfoil wear during contact with a surrounding casing.

Abstract: A rotor system for a gas turbine engine is provided having an outer rotor which includes an annular outer spool that comprises at least one ring portion of increased thickness with respect to the remainder of the outer spool, attached to a relatively thinner, axially extending arm portion. A plurality of outer turbine blades are attached to the ring portion of the outer spool and extend radially inwardly into a working gas flowpath of the engine. A casing surrounds the outer spool and defines an annular cavity between the casing and the outer spool, which is sealed by forward and aft brush seals at the forward and aft ends thereof. At least one inlet port is formed through the casing for admitting a flow of cooling air to the annular cavity. Means are provided for directing this cooling air flow to the ring portion of the outer spool.

Abstract: A blade retention assembly is provided that comprises: a disk, a plurality of blades, each having a dovetail disposed in one of an array of slots formed in the outer periphery of the disk, an annular inner retainer attached to the disk, and a plurality of arcuate blade retainers, the blade retainers being secured to the inner retainer by a hooked rabbet joint. A blade retainer is also provided, comprising an arcuate body having a planar inner face having an arcuate depression formed therein defining a sealing lip, and an aft face disposed opposite the inner face which has an arcuate rim formed therein.

Abstract: A rotor system for a gas turbine engine is provided having an outer rotor which includes an annular outer spool that comprises at least one ring portion of increased thickness with respect to the remainder of the outer spool, attached to a relatively thinner, axially extending arm portion. A plurality of outer turbine blades are attached to the ring portion of the outer spool and extend radially inwardly into a working gas flowpath of the engine. A casing surrounds the outer spool and defines an annular cavity between the casing and the outer spool, which is sealed by forward and aft brush seals at the forward and aft ends thereof. At least one inlet port is formed through the casing for admitting a flow of cooling air to the annular cavity. Means are provided for directing this cooling air flow to the ring portion of the outer spool.

Abstract: A blade retention assembly is provided that comprises: a disk, a plurality of blades, each having a dovetail disposed in one of an array of slots formed in the outer periphery of the disk, an annular inner retainer attached to the disk, and a plurality of arcuate blade retainers, the blade retainers being secured to the inner retainer by a hooked rabbet joint. A blade retainer is also provided, comprising an arcuate body having a planar inner face having an arcuate depression formed therein defining a sealing lip, and an aft face disposed opposite the inner face which has an arcuate rim formed therein.