Abstract: A propulsion system for an aircraft includes an electric propulsion engine configured to be mounted to the aircraft at an aft end of the aircraft. The electric propulsion engine includes a power gearbox mechanically coupled to an electric motor. The electric propulsion engine further includes a fan rotatable about a central axis of the electric propulsion engine by the electric motor through the power gearbox. Moreover, the electric propulsion engine includes an attachment assembly for mounting at least one of the electric motor or the power gearbox. The attachment assembly includes a torsional damper for accommodating a torsional vibration of the electric motor or the power gearbox.

Abstract: A propulsion system for an aircraft includes an electric propulsion engine configured to be mounted to the aircraft at an aft end of the aircraft. The electric propulsion engine includes a fan rotatable about a central axis of the electric propulsion engine. The fan includes a fan shaft mechanically coupled to a power gearbox. The electric propulsion engine also includes an electric motor having a driveshaft, with the electric motor coupled to the power gearbox through the drive shaft. The electric motor rotates the fan through the power gearbox. The driveshaft includes a flexible element for accommodating a misalignment of the electric motor and the power gearbox.

Abstract: A method and system for a turbofan gas turbine engine system is provided. The gas turbine engine system includes a variable pitch fan (VPF) assembly coupled to a first rotatable shaft and a low pressure compressor LPC coupled to a second rotatable shaft. The LPC including a plurality of variable pitch stator vanes interdigitated with rows of blades of a rotor of the LPC. The gas turbine engine system also includes a speed reduction device coupled to said first rotatable shaft and said second rotatable shaft. The gas turbine engine system further includes a modulating pressure relief valve positioned between an outlet of said LPC and a bypass duct and a controller configured to schedule a position of said plurality of variable pitch stator vanes and said modulating pressure relief valve in response to an operational state of said turbofan gas turbine engine system and a temperature associated with said LPC.

Abstract: A propulsion system for an aircraft includes an electric generator and a turbomachine. The turbomachine is configured to be mounted to a first wing of the aircraft and is operable with the electric generator. The propulsion system additionally includes a first propulsor mechanically coupled to a shaft of the turbomachine and a second propulsor assembly configured to be mounted at a location away from the turbomachine and the first propulsor. The electric generator is in electrical communication with the second propulsor assembly for powering the second propulsor assembly.

Abstract: A propulsion system for an aircraft includes an electric generator and a turbomachine. The turbomachine is configured to be mounted to a first wing of the aircraft and is operable with the electric generator. The propulsion system additionally includes a first propulsor mechanically coupled to a shaft of the turbomachine and a second propulsor assembly configured to be mounted at a location away from the turbomachine and the first propulsor. The electric generator is in electrical communication with the second propulsor assembly for powering the second propulsor assembly.

Abstract: A gas turbine engine includes a core and a tie shaft. The tie shaft is rotatable about an axial direction of the gas turbine engine by the core of the gas turbine engine. The gas turbine engine additionally includes a modular fan having a plurality of fan blades and a frame. The plurality fan blades are attached to the frame and the frame is slidably received on the tie shaft of the gas turbine engine. The modular fan additionally includes an attachment member removably attaching the frame to the tie shaft of the gas turbine engine to removably install the modular fan in the gas turbine engine.

Abstract: A planet gear train for an engine casing includes a sun gear rotatable by a shaft, a ring gear, a plurality of planet gears rotatably mounted in a planet carrier and meshing with the sun gear and the ring gear, and a damping system. The damping system includes a soft mount disposed about the shaft and coupling the planet carrier the engine casing. The soft mount includes a flexible inner sleeve and a flexible outer sleeve. The flexible inner sleeve and the flexible outer sleeve are fixedly coupled together at respective central portions thereof near the shaft. The damping system further includes a plurality of damping couplers disposed about an outer flange of the planet carrier. The plurality of damping couplers is configured to flexibly couple the flexible inner sleeve to the flexible outer sleeve.

Abstract: A method of manufacturing a frangible laminate is provided. The method comprises constructing a reinforced polymer matrix, cutting the reinforced polymer matrix into a plurality of laminae, forming a laminate via stacking the plurality of laminae and at least one energy dissipating member, and consolidating the laminate.

Abstract: A frangible airfoil that mitigates adverse conditions associated with release of material resulting from impact damage to the composite blade is provided, the airfoil having provisions for dissipating energy, self-shredding, and predetermined release trajectory.

Abstract: A method and system for a turbofan gas turbine engine system is provided. The gas turbine engine system includes a variable pitch fan (VPF) assembly coupled to a first rotatable shaft and a low pressure compressor LPC coupled to a second rotatable shaft. The LPC including a plurality of variable pitch stator vanes interdigitated with rows of blades of a rotor of the LPC. The gas turbine engine system also includes a speed reduction device coupled to said first rotatable shaft and said second rotatable shaft. The gas turbine engine system further includes a modulating pressure relief valve positioned between an outlet of said LPC and a bypass duct and a controller configured to schedule a position of said plurality of variable pitch stator vanes and said modulating pressure relief valve in response to an operational state of said turbofan gas turbine engine system and a temperature associated with said LPC.

Abstract: A rotary machine with at least one frangible composite blade that mitigates adverse conditions associated with release of material resulting from impact damage to the composite blade is provided. The composite blade having provisions for dissipating energy, self-shredding, and predetermined release trajectory. A method for manufacturing the composite blade and assembling the blade into a rotary machine is also provided.

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: A gas turbine propulsion system includes a system which utilizes a cryogenic liquid fuel for a non-combustion function.

Abstract: A cryogenic fuel storage system for an aircraft is disclosed including a cryogenic fuel tank having a first wall forming a storage volume capable of storing a cryogenic liquid fuel; an inflow system capable of flowing the cryogenic liquid fuel into the storage volume; an outflow system adapted to deliver the cryogenic liquid fuel from the cryogenic fuel storage system; and a vent system capable of removing at least a portion of a gaseous fuel formed from the cryogenic liquid fuel in the storage volume.

Abstract: An aircraft system comprising a fuel storage system comprising a first fuel tank capable of storing a first fuel and a second fuel tank capable of storing a second fuel is disclosed herein. The aircraft system further comprises a fuel cell system comprising a fuel cell capable of producing electrical power using at least one of the first fuel or the second fuel, and a fuel delivery system capable of delivering a fuel from the fuel storage system to the fuel cell system.

Abstract: A support structure for a gas turbine engine includes a support arm (214, 216) extending between a low pressure shaft (202) and a rotor. The support arm includes a fuse (220) having a low failure point. A high pressure stub shaft (270) axially and radially supports the low pressure shaft (202) after fuse failure. An axial gap (A) between a portion of the low pressure shaft and the stub shaft permits movement of the low pressure shaft after fuse failure. A radial gap (B) between the stub shaft (270) and the low pressure shaft (202) allows radial deflection of the low pressure rotor system after fuse failure.