Abstract: A gas turbine engine is provided including a turbine section including a turbine having a plurality of first speed turbine rotor blades; a compressor section including a compressor having a plurality of first speed compressor rotor blades and a plurality of second speed compressor rotor blades; a gearbox; and a first spool rotatable by the plurality of first speed turbine rotor blades, the first spool coupled to the plurality of first speed compressor rotor blades for driving the plurality of first speed compressor rotor blades in a first direction and to the plurality of second speed compressor rotor blades across the gearbox for driving the plurality of second speed compressor rotor blades in a second direction, opposite the first direction.

Abstract: An aircraft includes a fuselage; a wing coupled to, and extending from, the fuselage; and a propulsion system. The propulsion system includes a plurality of electric fans integrated into the wing and oriented to generate thrust along a vertical direction, the plurality of electric fans arranged along a length of the wing and including an outer-most electric fan along a transverse direction relative to the fuselage. The outer-most electric fan is at least one of a variable pitch fan or a variable speed fan to provide increased stability to the aircraft.

Abstract: An unducted thrust producing system has a rotating element with an axis of rotation and a stationary element. The rotating element includes a plurality of blades, and the stationary element has a plurality of vanes configured to impart a change in tangential velocity of the working fluid opposite to that imparted by the rotating element acted upon by the rotating element. The system includes an inlet forward of the rotating element and the stationary element.

Abstract: A shroud assembly for a gas turbine engine includes: at least one shroud hanger; an annular array of shroud segments carried by the child hanger at least one and arrayed about an engine centerline axis, each of the shroud segments having a flowpath side defining a portion of a primary engine flowpath, and an opposed backside facing the at least one shroud hanger, wherein the shroud segment is mounted to the at least one shroud hanger such that it is movable in a radial direction between an inboard position and an outboard position, in response to a balance of gas pressures prevailing on the flowpath side and the backside; and biasing means which urge each of the shroud segments towards one of the positions.

Abstract: A unshrouded vane assembly for an unducted propulsion system includes a plurality of vanes which have non-uniform characteristics configured to generate a desired vane exit swirl angle.

Abstract: A noise reducing airfoil defining a span extending between a root and a tip and a chord at each point along the span extending between a leading edge and a trailing edge. The airfoil includes a pressure side, a suction side, and a trailing edge sheath including an outer surface coupled to the trailing edge of the airfoil. The trailing edge sheath extends at least partially along the chord on the pressure and suction sides at each point along the span within the trailing edge sheath. The trailing edge sheath defines a fluid passageway extending along at least a portion of the span. Further, the trailing edge sheath defines at least one aperture on at least one of the pressure side, the suction side, or trailing edge fluidly coupling the fluid passageway to the outer surface.

Abstract: A lift fan apparatus includes: a rotor having at least one rotatable hub carrying at least one row of blades; a duct surrounding the rotor, the duct including spaced-apart peripheral walls extending between an inlet and an exit, the peripheral walls collectively defining a flow channel which includes a diffuser disposed downstream of the rotor, in which a flow area at the exit is greater than a flow area at the rotor; and a plurality of spaced-apart splitters disposed in the diffuser, each of the splitters having opposed side walls extending between an upstream leading edge and a downstream trailing edge, wherein the splitters divide the diffuser into a plurality of side-by-side flow passages.

Abstract: A fan for a gas turbine engine includes: an annular casing; a disk disposed inside the casing and mounted for rotation about an axial centerline, the disk including a row of fan blades extending radially outwardly therefrom; each of the fan blades including an airfoil having circumferentially opposite pressure and suction sides extending radially in span from a root to a tip, and extending axially in chord between spaced-apart leading and trailing edges, with the airfoils defining corresponding flow passages therebetween for pressurizing air; the row including no more than 21 and no less than 13 of the fan blades; and wherein each of the fan blades has a solidity defined by a ratio of the airfoil chord over a circumferential pitch of the fan blades, measured at 60% of a radial distance from the root to the tip, of less than about 1.6.

Abstract: A propulsion device that defines a central axis and a circumferential direction is provided. The propulsion device may include a core engine and a core casing. The core engine may include an engine shaft extending along the central axis. The core casing may have an inner surface and an outer surface. The core casing may extend along the circumferential direction about the propulsion device, as well as along the central axis from a forward end to an aft end. The core casing may define a primary air flowpath having an annular inlet at the forward end and an exhaust at the aft end. The core casing may further define a reverse flow passage extending from an outer surface entrance to an inner surface exit.

Abstract: A variable-solidity propeller apparatus, comprising a propeller having at least one rotatable hub carrying at least one row of propeller blades, wherein the propeller has a first configuration in which the propeller has a first average solidity, and a second configuration in which the propeller has a second average solidity which is greater than the first average solidity.

Abstract: A fan for a gas turbine engine includes: an annular casing; a disk disposed inside the casing and mounted for rotation about an axial centerline, the disk including a row of fan blades extending radially outwardly therefrom; each of the fan blades including an airfoil having circumferentially opposite pressure and suction sides extending radially in span from a root to a tip, and extending axially in chord between spaced-apart leading and trailing edges, with the airfoils defining corresponding flow passages therebetween for pressurizing air; the row including no more than 21 and no less than 13 of the fan blades; and wherein each of the fan blades has a solidity defined by a ratio of the airfoil chord over a circumferential pitch of the fan blades, measured at 60% of a radial distance from the root to the tip, of less than about 1.6.

Abstract: A containment structure comprising an annular structure configured to provide close tolerance to a rotation structure. The containment structure includes a first abradable structure, a second abradable structure, and a shroud that is defined in part by the first abradable structure and in part by the second abradable structure. An annular surface of the shroud is defined by the first abradable structure and by the second abradable structure. The first abradable structure is harder than the second abradable structure.

Abstract: An apparatus and method of controlling a multi-stage compressor of a gas turbine engine having a front-block of stages with variable stator vanes (VSVs), a rear block of stages downstream of the front-block, and a bleed air off-take from at least one of the stages. The method includes sensing the pressure of the bleed air from the bleed air off-take and adjusting the position of the VSV for the front-block.

Abstract: An unducted thrust producing system is provided that can include a rotating element having an axis of rotation about a central longitudinal axis and comprising a plurality of blades attached to a spinner; a stationary element positioned aft of the rotating element; and an inlet positioned between the rotating element and the stationary element such that the inlet passes radially inward of the stationary element. The rotating element defines an annular valley positioned between a first annular crown and a second annular crown, and the inlet defines an open area positioned aft of the second annular crown.

Abstract: A heat exchanger apparatus includes: spaced-apart peripheral walls extending between an inlet and an outlet, the peripheral walls collectively defining a flow channel which includes a diverging portion downstream of the inlet, in which a flow area is greater than a flow area at the inlet; a plurality of spaced-apart fins disposed in the flow channel, each of the fins having opposed side walls extending between an upstream leading edge and a downstream trailing edge, wherein the fins divide at least the diverging portion of the flow channel into a plurality of side-by-side flow passages; and a heat transfer structure disposed within at least one of the fins.

Abstract: A variable pitch fan for a propulsion device is provided. The variable pitch fan includes a plurality of fan blades coupled to a disk and a rotatable front hub covering the disk. Each fan blade is rotatable about a pitch axis to vary a pitch of the fan blade. The pitch is variable within a pitch range that is at least about 80° to about 130°. Each fan blade extends radially outward from the disk along a span from a root to a tip. A portion of the span adjacent the root defines a root span region, a portion of the span adjacent the tip defines a tip span region. A solidity of the variable pitch fan is at least 1.0 in the root span region.

Abstract: An unducted thrust producing system has a rotating element with an axis of rotation and a stationary element. The rotating element includes a plurality of blades, each having a blade root proximal to the axis, a blade tip remote from the axis, and a blade span measured between the blade root and the blade tip. The rotating element has a load distribution such that at any location between the blade root and 30% span the value of ?RCu in the air stream is greater than or equal to 60% of the peak ?RCu in the air stream.

Abstract: A gas turbine engine comprises a fan, a compressor, a turbine having first stage blades and second stage blades, the first stage blades rotates in a first direction and the second stage blades rotates in a second direction opposite the first direction. The first stage blades are directly adjacent to one another, a drive in operable input connection with the fan and in operable output connection with the first stage blades and the second stage blades, the first stage blades and the second stage blades driving the fan through the drive.

Abstract: An aircraft includes a fuselage; a propulsion system including a power source and a vertical thrust propulsor driven by the power source; and a wing extending from the fuselage, the vertical thrust propulsor positioned on or at least partially within the wing, the wing including a diffusion assembly, the diffusion assembly including at least one diffusion member fixed in position and located downstream of the vertical thrust propulsor for diffusing an airflow from the vertical thrust propulsor.

Abstract: An aircraft includes a fuselage; a propulsion system including a power source and a plurality of vertical thrust electric fans driven by the power source; and a wing extending from the fuselage. The plurality of vertical thrust electric fans are arranged along the wing, the wing including one or more components being moveable to selectively expose the plurality of vertical thrust electric fans, the wing further including a diffusion assembly positioned downstream of at least one vertical thrust electric fan of the plurality of vertical thrust electric fans, the diffusion assembly defining a diffusion area ratio greater than 1:1 and less than about 2:1.