Abstract: A structure for damping at a fluid manifold assembly for an engine is generally provided. The fluid manifold assembly includes a first walled conduit defining a first fluid passage therewithin. A flow of fluid defining a first frequency is permitted through the first fluid passage. A second walled conduit includes a pair of first portions each coupled to the first walled conduit. A second portion is coupled to the pair of first portions. A second fluid passage is defined through the first portion and the second portion in fluid communication with the first fluid passage. The flow of fluid is permitted through the second fluid passage at a second frequency approximately 180 degrees out of phase from the first frequency.

Abstract: In one an exemplary aspect of the present disclosure, an engine includes a drive shaft; an electric machine including a stator assembly and a rotor assembly, the rotor assembly rotatable relative to the stator assembly; and an electrical break, the drive shaft coupled to the rotor assembly through the electrical break.

Abstract: A gas turbine engine includes a fan section and a core turbine engine operable with the fan section. The gas turbine engine also includes a sump pressurization assembly including a valve, a first duct in airflow communication with the valve and a first pressurized air source, and a second duct in airflow communication with the valve and a second pressurized air source. Further, the sump pressurization assembly includes a supply duct selectively in airflow communication with the first duct and the second duct through the valve. The supply duct is located at least partially inward of a core air flowpath of the core turbine engine for pressurizing a sump of the gas turbine engine.

Abstract: The present disclosure is directed to a gas turbine engine defining a radial direction, a circumferential direction, an axial centerline along a longitudinal direction, and wherein the gas turbine engine defines an upstream end and a downstream end along the longitudinal direction. The gas turbine engine includes a first turbine rotor comprising an inner shroud, an outer shroud outward of the inner shroud in the radial direction, at least one connecting airfoil coupling the inner shroud and the outer shroud at least partially along the radial direction, and an outer band outward of the outer shroud in the radial direction and extended at least partially in the circumferential direction, and a plurality of connecting members couples the outer shroud and the outer band.

Abstract: A gas turbine engine including a core having in serial flow order a compressor, a combustor, and a turbine—the compressor, combustor, and turbine together defining a core air flowpath. The gas turbine engine additionally includes a fan section mechanically coupled to the core, the fan section including a plurality of fan blades, and each of the plurality fan blades defining a pitch axis. An actuation device is operable with the plurality fan blades for rotating the plurality fan blades about their respective pitch axes, the actuation device including an actuator located outward of the core air flowpath to, e.g., simplify the gas turbine engine.

Abstract: A self-contained lubrication fluid circulation system for use with a gas turbine engine defining a core air flowpath includes a sump configured to collect lubrication fluid and a heat source coupled in fluid communication with the sump and configured to transfer heat to the lubrication fluid. The system also includes a heat sink positioned within the sump and coupled in fluid communication with the heat source. A lubrication fluid conduit of the system is configured to channel the lubrication fluid between the heat source and the heat sink, wherein the lubrication fluid conduit is positioned entirely within the sump.

Abstract: A method and system for controlling a pitch of blades of a fan assembly having a centerline axis of rotation is provided. The system includes a pitch change mechanism (PCM) including a hydraulic actuator positioned axisymmetric with respect to the fan assembly and configured to angularly displace the blades of the fan assembly between a first position and a second position. The PCM further includes a plurality of hydraulic fluid supply lines coupled in flow communication between the hydraulic actuator and a hydraulic fluid transfer sleeve, the hydraulic fluid transfer sleeve configured to transfer a flow of pressurized hydraulic fluid across a gap between a stationary member of the hydraulic fluid transfer sleeve and a rotatable member of the hydraulic fluid transfer sleeve.

Abstract: The present disclosure is directed to a gas turbine engine defining a radial direction, a circumferential direction, an axial centerline along a longitudinal direction. The gas turbine engine defines an upstream end and a downstream end along the longitudinal direction and includes a turbine frame defined around the axial centerline. The turbine frame includes a first bearing surface, a second bearing surface, and a third bearing surface. The first bearing surface corresponds to a first turbine rotor, the second bearing surface corresponds to a second turbine rotor, and the third bearing surface corresponds to a third turbine rotor, and each turbine rotor is independently rotatable.

Abstract: A system for heat exchange and leak detection is generally provided, the system including a heat exchanger including a first wall defining a first passage containing a first fluid. A leak detection enclosure containing a leak detection medium is defined between the first wall and a second wall surrounding the first wall.

Abstract: A gas turbine engine including a fan, a core including at least one rotatable shaft, and a gearbox mechanically coupling at least one rotatable shaft of the core to the fan is provided. The gas turbine engine also includes a coupling system for mounting the gearbox within the gas turbine engine. The coupling system includes a flexible coupling connected to at least one of a fan frame or a core frame, as well as a torque frame connected to the flexible coupling and the gearbox. Moreover, a deflection limiter is provided, loosely attaching the flexible coupling to the gearbox to provide a predetermined axial range of motion, radial range of motion, and circumferential range of motion between the gearbox and the frame to which the flexible coupling is attached.

Abstract: A gas turbine engine including a compressor section, a turbine section, and a combustion section positioned between the compressor section and the turbine section is provided. The gas turbine engine also includes a cooling system having a tank and one or more fluid lines in fluid communication with the tank. The one or more fluid lines are configured to carry a flow of consumable cooling liquid provide such consumable cooling liquid to one or more components of the compressor section, the turbine section, and/or the combustion section not directly exposed to a core air flowpath defined through the gas turbine engine.

Abstract: A gas turbine engine includes a compressor section and a turbine section together defining a core air flowpath. The rotary component is rotatable with the compressor section and the turbine section. The gas turbine engine additionally includes an electric machine rotatable with the rotary component and positioned coaxially with the rotary component at least partially inward of the core air flowpath. The electric machine is flexibly mounted to a static frame member, or flexibly coupled to the rotary component, or both, such that the electric machine is mechanically isolated or insulated from various internal and external forces on the gas turbine engine.

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: An annular duct including a modular annular heat exchanger for a gas turbine engine is provided, where the modular annular heat exchanger includes a plurality of radial modules in circumferentially adjacent arrangement. Each radial module includes a cooled fluid inlet plenum segment, a plurality of blades, and a cooled fluid outlet plenum segment. The plurality of blades is configured in circumferentially adjacent arrangement and defines an angular space that is conformal between each circumferentially adjacent blade. The cooled fluid inlet plenum segment, the plurality of blades, and the cooled fluid outlet plenum segment are in serial axial flow arrangement and define an internal cooled fluid flowpath and an external cooling fluid flowpath parallel to the internal cooled fluid flowpath. Each radial module further includes an inner annular ring segment and an outer annular ring segment. The inner annular ring segment and the outer annular ring segment define a plurality of blade retainers.

Abstract: Systems and methods for estimating engine and electrical machine health are provided. The systems and methods therefore include features that provide for improved health estimates of one or more engines and/or electrical machines of one or more engines having an electrical machine embedded therein or coupled thereto. In particular, the systems and methods therefore include features for utilizing sensed or measured operating parameters of the electrical machines and/or operating parameters of the prime mover to estimate the health of the prime mover and/or the electrical machines embedded therein or coupled thereto.

Abstract: A propulsion system for an aircraft includes at least two propulsion engines. Each propulsion engine rotates about a longitudinal rotational axis, and a propulsive fan associated with each propulsion engine rotates in a plane perpendicular to the longitudinal rotational axis. Each plane is at least one of spaced axially along a fuselage of the aircraft and canted such that each plane does not intersect any other plane associated with another propulsion engine and propulsive fan.

Abstract: A variable pitch fan for gas turbine engine is provided. The variable pitch fan includes a plurality of fan blades rotatably coupled to a disk and an actuation assembly for changing a pitch of each of the plurality of fan blades. The actuation assembly generally includes an actuation member operably connected to at least one of the plurality of fan blades and a pin. The pin is movable between a first position in which the pin is positioned at least partially in a channel defined in the actuation member, and a second position. The pin blocks movement of the actuation member relative to the pin past a range defined by the channel when in the first position. The actuation assembly further includes a retraction system for selectively engaging the pin in moving the pin from the first position to the second position.

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 propulsion system for an aircraft is provided having a propulsion engine configured to be mounted to the aircraft. The propulsion engine includes an electric machine defining an electric machine tip speed during operation. The propulsion system additionally includes a fan rotatable about a central axis of the electric propulsion engine with the electric machine. The fan defines a fan pressure ratio, RFP, and includes a plurality of fan blades, each fan blade defining a fan blade tip speed. The electric propulsion engine defines a ratio of the fan blade tip speed to electric machine tip speed that is within twenty percent of the equation, 1.68×RFP?0.518, such that the propulsion engine may operate at a desired efficiency.

Abstract: A shaft assembly for a rotary machine, the rotary machine includes a low pressure compressor, a high pressure compressor, a high pressure turbine, and a low pressure turbine. The shaft assembly includes an outer shaft configured to couple the high pressure compressor to the high pressure turbine. The outer shaft includes a first forward bearing assembly positioned forward of the high pressure compressor, and a first aft bearing assembly positioned between the high pressure compressor and the high pressure turbine. The shaft assembly also includes an inner shaft rotatable about a common axis with the outer shaft that at least partially extends about the inner shaft. The inner shaft is configured to couple the low pressure compressor to the low pressure turbine. The inner shaft includes a second forward bearing assembly positioned forward of the low pressure compressor, and a second aft bearing assembly positioned proximate the high pressure turbine.