Abstract: A fan section for a gas turbine engine is disclosed herein. The fan may include a rotor disk and a plurality of airfoils fixedly attached to and supported by the rotor disk as a single unitary piece. The airfoils may extend radially outward from the rotor disk with respect to an engine axis. The rotor disk may be made of metal and the airfoils may each be made at least partially of an organic matrix composite.

Abstract: A gas turbine engine includes a fan section and a speed change mechanism for driving the fan section. The speed change mechanism is an epicyclic gear train. A torque frame surrounds the speed change mechanism and includes a plurality of fingers. A bearing support is attached to the plurality of fingers. A first fan section support bearing is mounted forward of the speed change mechanism and a second fan section bearing is mounted on the bearing support aft of the speed change mechanism. The second fan section bearing is a fan thrust bearing.

Abstract: A gas turbine engine includes a fan, a compressor section, a combustor, and a turbine section where the turbine section is downstream of the combustor section. A shaft connects the turbine section to the compressor section. A bore tube is disposed within the shaft downstream of the compressor section. The bore tube includes an inlet connected to an air source for passing cooling air in an upstream direction of the shaft.

Abstract: A heat exchanger duct includes a wall having ends spaced along a central axis. An inlet manifold is positioned within a downstream portion of the duct at a radially outward location. An outlet manifold is positioned within an upstream portion of the duct at a radially outward location. At least one of the inlet and outlet manifolds extend at least 10 degrees around the circumference of the duct. A central manifold is disposed between the inlet and outlet manifolds, and radially inwardly of the inlet and outlet manifolds. Heat exchanger entrance elements extend radially inward from the inlet manifold to the central manifold, and heat exchanger exit elements extend radially outward from the central manifold to the outlet manifold. A gas turbine engine is also disclosed.

Abstract: A gas turbine engine includes a fan rotor, a compressor section, a combustor section, and a turbine section. The turbine section is positioned downstream of the combustor section. A fan drive turbine in the turbine section, and a shaft connects the fan drive turbine to the fan rotor. An inlet duct is connected to a cooling air source and connected to a cooling compressor downstream of the fan drive turbine. The cooling compressor is connected to an air source, and connected to a turning duct for passing compressed air in an upstream direction through the shaft. A method is also disclosed.

Abstract: A rotor assembly is provided for a gas turbine engine. This rotor assembly includes a first rotor disk, a second rotor disk, a plurality of rotor blades and a plurality of disk mounts. The first rotor disk is configured to rotate about a rotational axis. The second rotor disk is configured to rotate about the rotational axis. The rotor blades are arranged circumferentially around the rotational axis. Each of the rotor blades is mounted to the first rotor disk and to the second rotor disk. The rotor blades include a first rotor blade. Each of the disk mounts connects the first rotor disk and the second rotor disk together. The disk mounts include a first disk mount that further supports the first rotor blade.

Abstract: An apparatus is provided for a rotor assembly with a rotational axis. This apparatus includes a rotor blade including an airfoil, a neck and an attachment. The neck radially connects the airfoil and the attachment. The attachment extends axially between an attachment first axial side and an attachment second axial side. A first end portion of the attachment projects axially out from the neck to the attachment first axial side. A second end portion of the attachment projects axially out from the neck to the attachment second axial side.

Abstract: A rotor assembly is provided for a gas turbine engine. This rotor assembly includes a first rotor disk, a second rotor disk, a plurality of rotor blades and a plurality of disk mounts. The first rotor disk is configured to rotate about a rotational axis. The second rotor disk is configured to rotate about the rotational axis. The rotor blades are arranged circumferentially around the rotational axis. Each of the rotor blades is axially between and mounted to the first rotor disk and the second rotor disk. The disk mounts connect the first rotor disk and the second rotor disk together. The disk mounts include a first disk mount. The first disk mount is integral with the first rotor disk. The first disk mount projects axially through the second rotor disk.

Abstract: A rotor blade is provided for a gas turbine engine. This rotor blade includes a rotor blade pair including a mount, a first airfoil and a second airfoil. The mount includes a forked body with a first leg and a second leg. The first airfoil is connected to the first leg. The second airfoil is connected to the second leg and arranged circumferentially next to the first airfoil.

Abstract: A rotor assembly is provided for a gas turbine engine. This rotor assembly includes a first rotor disk, a second rotor disk, a plurality of rotor blades and a plurality of vanes. The first rotor disk is configured to rotate about a rotational axis. The first rotor disk is configured from or otherwise includes disk material. The second rotor disk is configured to rotate about the rotational axis. The rotor blades are arranged circumferentially around the rotational axis. Each of the rotor blades is axially between and mounted to the first rotor disk and the second rotor disk. The vanes are arranged circumferentially around the rotational axis and axially between the first rotor disk and the second rotor disk. The vanes include a first vane, which first vane is configured from or otherwise includes vane material that is different than the disk material.

Abstract: A gas turbine engine includes a turbine rotor connected to a main compressor rotor. A tail cone is mounted inward of an exhaust core flow. A generator rotor is adjacent a generator stator. The generator rotor and stator are mounted within the tail cone. A passage connects a bypass flow path to the tail cone. A cooling air compressor is operable within the passage. The turbine rotor drives a shaft to drive the generator rotor and the cooling compressor. A method is also disclosed.

Abstract: A rotor assembly is provided for a gas turbine engine. This rotor assembly includes a first rotor disk, a second rotor disk and a plurality of rotor blades. The first rotor disk is configured to rotate about a rotational axis. The second rotor disk is configured to rotate about the rotational axis. The rotor blades are arranged circumferentially around the rotational axis. Each of the rotor blades is mounted to the first rotor disk and to the second rotor disk. The rotor blades include a first rotor blade. The first rotor blade includes an attachment projecting axially along the rotational axis into a first pocket in the first rotor disk and a second pocket in the second rotor disk. The attachment has a dovetail cross-sectional geometry when viewed in a plane perpendicular to the rotational axis. A portion of the first rotor disk extends circumferentially across and covers the attachment.

Abstract: A rotor assembly is provided for a gas turbine engine. This rotor assembly includes a first rotor disk, a second rotor disk, a plurality of rotor blades and a plurality of vanes. The first rotor disk is configured to rotate about a rotational axis. The second rotor disk is configured to rotate about the rotational axis. The rotor blades are arranged circumferentially around the rotational axis. Each of the rotor blades is axially between and mounted to the first rotor disk and the second rotor disk. The vanes are arranged circumferentially around the rotational axis. The vanes include a first vane that is integral with the first rotor disk and projects axially to the second rotor disk.

Abstract: A cooling system for a gas turbine engine may comprise a plenum extending circumferentially around an outer engine case structure. The plenum may comprise a supply conduit and a return conduit. The supply conduit and the return conduit may be in fluid communication with a heat exchanger. The heat exchanger may be disposed between the outer engine case structure and an inner engine case structure. The plenum may be configured to provide enhance heat transfer for the cooling system.

Abstract: A fuel power transfer system for an engine may include a cryogenic fuel supply, a fuel pump in fluid communication with the cryogenic fuel supply, a multi-position valve in fluid communication with the fuel pump and a combustion chamber of the engine, a fuel turbine operatively coupled to the fuel pump and having a primary discharge port in fluid communication with the combustion chamber, a primary heat exchanger in fluid communication between the multi-position valve and the fuel turbine, and a gearbox operatively coupled to the fuel turbine and the fuel pump and configured to transfer power from the fuel turbine to the engine.

Abstract: A cryogenic fuel auxiliary power system for an engine may include a cryogenic fuel supply, a first valve in fluid communication with the cryogenic fuel supply and configured to control a fuel flow, a first heat exchanger, configured to receive the fuel flow, in fluid communication with the first valve and a combustion chamber of the engine, and a fuel cell in fluid communication between the first valve and the first heat exchanger.

Abstract: A gas turbine engine system includes a gas turbine engine and a fuel turbine system. The gas turbine engine includes a heat exchange system configured to transfer thermal energy from a first compressed air flow and an exhaust gas flow to a fuel to produce a gaseous fuel. The fuel turbine system includes a fuel turbine fluidly coupled to the heat exchange system and a combustor of the gas turbine engine, and a fuel pump fluidly coupled to the heat exchange system and configured to be driven by the fuel turbine. The fuel turbine is configured to extract energy from expansion of the gaseous fuel to produce the gaseous fuel at a lower pressure for delivery to the combustor.

Abstract: An aircraft power system according to an exemplary embodiment of this disclosure includes, among other possible things, a battery, a motor/generator coupled to the battery, an inertial drum rotatable about an axis of rotation and coupled to the motor/generator, wherein the motor/generator drives rotation of the inertial drum in a first operating mode and is driven by the inertial drum in a second operating mode; and a housing defining a chamber for the inertial drum, the chamber filled with low-viscosity medium to reduce friction on the inertial drum.

Abstract: A gas turbine engine system includes a gas turbine engine and a turbo-generator. The gas turbine engine includes a heat exchange system configured to transfer thermal energy from an air flow (i.e., inlet air flow or exhaust gas flow) to a fuel to produce a gaseous fuel. The turbo-generator includes a fuel turbine fluidly coupled to the heat exchange system and a combustor of the gas turbine engine, a fuel pump configured to be driven by the fuel turbine and fluidly coupled to the heat exchange system, and a motor/generator configured to be driven by the fuel turbine. The fuel turbine is configured to extract energy from expansion of the gaseous fuel to produce a gaseous fuel for combustion in the combustor. The motor/generator includes a cooling jacket, which is fluidly coupled to the fuel pump.

Abstract: A light weight component, the light weight component including: a metallic foam core formed into a desired configuration; an external metallic shell applied to an exterior surface of the metallic foam core after it has been formed into the desired configuration; an inlet opening and an outlet opening formed in the external metallic shell in order to provide a fluid path through the metallic foam core; and a thermoplastic material injected into the metallic foam core via the inlet opening.