Abstract: A gas turbine engine section includes a plurality of spaced rotor stages, with a static guide vane intermediate the spaced rotor stages. The static guide vane provides swirl into air passing toward a downstream one of the spaced rotor stages, and an outer housing surrounding the spaced rotor stages. A diverter diverts a portion of air radially outwardly through the outer housing, and across at least one heat exchanger. The diverted air passes back into a duct radially inwardly through the outer housing, and is exhausted toward the downstream one of the spaced rotor stages.

Abstract: An example method of cooling a compressor of a gas turbine includes, among other things, diverting a flow from a compressor, and directing the flow at the compressor in a direction, the direction having a circumferential component and an axial component.

Abstract: A structurally-reinforced rotor disk for a gas turbine engine is disclosed. The rotor disk may comprise a body including a rim configured to support airfoils (which may be separate or integral with the airfoils), an axially-extending bore disposed radially inward of the rim, and a radially-extending web connecting the rim and the bore. The bore may include an axial outer edge and at least one circumferentially-extending annular recess formed axially between the outer edge and the web. The rotor disk may further comprise an annular ring retained in the annular recess, and the annular ring may be formed from a different material than the body of the rotor disk so as to increase a self-sustaining radius of the rotor disk.

Abstract: According to one aspect of the present disclosure, a gas turbine engine is disclosed that includes an engine section comprising a plurality of stages of variable vanes, and also includes first and second synchronizing rings (sync-rings). Movement of the first sync-ring adjusts vane angles of a first one of the stages of variable vanes, and movement of the second sync-ring adjusts vane angles of a second one of the stages of variable vanes. At least one sensor is configured to measure a condition of the gas turbine engine. A controller is configured to move the first sync-ring independently of the second sync-ring based on data from the at least one sensor.

Abstract: Disclosed is an assembly for a gas turbine engine, the assembly includes: a spinner or nosecone comprising a threaded rear portion, an engine structure comprising a threaded front portion, the nosecone being threadingly connected to the engine structure, wherein rotation of the spinner or nosecone about the engine structure in a first direction secures the spinner or nosecone to the engine structure and rotation of the spinner or nosecone about the engine structure in a second direction releases the spinner or nosecone from the engine structure; and a lock ring slidingly connected to the engine structure to slide between: a forward position to engage the spinner or nosecone and block rotation of the spinner or nosecone in the second direction, and a rearward position, where the lock ring is spaced from the spinner or nosecone.

Abstract: Aspects of the disclosure are directed to a gas turbine engine shaft, having an outer circumferential surface that circumscribes a central axis. The gas turbine engine shaft may comprise a plurality of splines extending along a portion of the outer circumferential surface, where each of the plurality of splines comprises an axial face that tapers from the outer circumferential surface to a spline top surface and includes a curved chamfered surface at a radially distal end of the axial face. The plurality of splines may further comprise a first sidewall and a second sidewall separated by a first spline width distance along the outer circumferential surface, where the first and second sidewalls taper inwardly from the outer circumferential surface to a second spline width along the spline top surface where the second spline width distance is less than the first spline width distance.

Abstract: A bleed valve system for a gas turbine engine that includes a retaining ring and a metering ring. The retaining ring abuts an end of a bleed air duct and defines a plurality of retaining ring apertures. The metering ring that is at least partially disposed within the bleed air duct and abuts the retaining ring, the metering ring defining a plurality of metering ring apertures, the metering ring being rotatable relative to the retaining ring to selectively facilitate a fluid flow through the bleed air duct.

Abstract: A lock for securing a nosecone having first threading to an inlet case having second threading and for use with a gas turbine engine having an axis includes an annular ring designed to be positioned about the first threading. The lock also includes a tab extending radially from the annular ring, and having a ramped circumferential end and an angled circumferential end such that the ramped circumferential end passes over a detent of the inlet case in response to the nosecone being fastened to the inlet case and the angled circumferential end engages with the detent to resist unfastening of the nosecone from the inlet case.

Abstract: A disclosed stator vane assembly includes a plurality of stator vane elements supported within a case about an axis. Each of the plurality of stator vane elements includes a platform having a recess on a first side and a tongue on a second side. The tongue overlaps a recess of an adjacent one of the plurality of stator vane elements with at least one of the recess and the tongue includes a crowned surface forming a seal.

Abstract: A variable vane system includes a harmonic drive driven by an actuator, and a drive gear driven by the harmonic drive. A geared unison ring includes an actuator gear rack driven by the drive gear and a multiple of vane gears, each of the multiple of vane gears mounted to one of a multiple of variable vanes, the multiple of vane gears driven by a gear rack of the geared unison ring.

Abstract: A turbine engine is disclosed and includes an airflow passage including an inner surface defined by a main shroud and a shroud extension. A flow splitter is disposed radially outward of the inner surface and axially overlapping the shroud extension. The turbine engine further includes a rotor including a blade proximate the shroud extension and an annular gap defined between the shroud extension and the blade of a first axial length less than a second axial length between the blade and an end of the flow splitter.

Abstract: A vane assembly for a gas turbine includes a first shroud segment, a second shroud segment, a first inner air seal, and a seal assembly. The first shroud segment has a first end face. The second shroud segment is disposed adjacent to the first shroud segment and has a second end face that faces towards and is spaced apart from the first end face by a gap. The first inner air seal extends into the first shroud segment and defines a first trench. The seal assembly at least partially received within the first trench.

Abstract: A gas turbine engine section includes a plurality of spaced rotor stages, with a static guide vane intermediate the spaced rotor stages. The static guide vane provides swirl into air passing toward a downstream one of the spaced rotor stages, and an outer housing surrounding the spaced rotor stages. A diverter diverts a portion of air radially outwardly through the outer housing, and across at least one heat exchanger. The diverted air passes back into a duct radially inwardly through the outer housing, and is exhausted toward the downstream one of the spaced rotor stages.

Abstract: A non-metallic engine case inlet compression seal for a gas turbine engine includes a non-metallic longitudinal leg section that extends from the non-metallic arcuate interface section and a non-metallic mount flange section that extends from the longitudinal leg section.

Abstract: A non-metallic engine case inlet compression seal for a gas turbine engine includes a non-metallic longitudinal leg section that extends from the non-metallic arcuate interface section and a non-metallic mount flange section that extends from the longitudinal leg section.

Abstract: A propulsive engine rotor configured for operation within a gas turbine includes a plurality of spaced airfoil blades circumferentially affixed to an outer hoop-style rim. At least two of the airfoil blades have an integral root configured to extend either through or at least radially inwardly from the rim. A woven fiber system provides that the blade roots are interdigitally wrapped by a plurality of woven fibers before impregnation and/or encapsulation of the fibers and wrapped roots within a ceramic or other composite matrix material to form a composite ring. The composite ring defines the interior body of the rotor, includes a bore through which passes the rotational axis of the rotor, and has a lower mass than the rotor rim. The woven fiber interface between the roots increases the tensile load capacity of the airfoil blades relative to the composite ring, and increases the self-sustaining radius of the rotor.

Abstract: A system for retaining stators and reducing air leakage in a gas turbine engine having an axis includes a stator having an inner platform, an outer platform, a low pressure side, a high pressure side, and at least one foot, and designed to turn air. The system also includes a case positioned radially outward from the stator and having at least one recess designed to interface with the at least one foot to resist movement of the stator relative to the case. The system also includes a bladder positioned between the outer platform of the stator and the case and designed to receive pressurized fluid having a greater pressure than ambient pressures experienced at the low pressure side of the stator and to further resist movement of the stator relative to the case in response to receiving the pressurized fluid.

Abstract: A lock for securing a nosecone having first threading to an inlet case having second threading and for use with a gas turbine engine having an axis includes an annular ring designed to be positioned about the first threading. The lock also includes a tab extending radially from the annular ring, and having a ramped circumferential end and an angled circumferential end such that the ramped circumferential end passes over a detent of the inlet case in response to the nosecone being fastened to the inlet case and the angled circumferential end engages with the detent to resist unfastening of the nosecone from the inlet case.

Abstract: A cooling system for providing a buffer cooled cooling air to a turbine section of a gas turbine engine is disclosed. The cooling system may comprise a first conduit configured to transmit a cooling air toward the turbine section, a heat exchanger configured to cool a bleed airflow diverted from the first conduit to provide a buffer air, and a bypass conduit configured to direct at least a portion of the buffer air through at least one passageway that bypasses a bearing compartment of the gas turbine engine. The cooling system may further comprise a manifold configured to allow the cooling air exiting the first conduit and the buffer air exiting the bypass conduit to mix and provide the buffer cooled cooling air, and a nozzle assembly configured to deliver the buffer cooled cooling air to the turbine section.

Abstract: A nose cone assembly for a gas turbine engine and method of circulating air in a gas turbine engine are disclosed. The nose cone assembly includes a nose cone having an aperture communicating air to an interior space of the nose cone and a discharge member communicating the air out of the nose cone.