Abstract: Dual bypass gas turbine engines and associated methods are provides. A dual bypass turbofan gas turbine engine includes counter-rotating first and second fans through which ambient air is received and propelled in an aft direction; a first bypass duct receiving first bypass air including a first portion of the propelled air; a second bypass duct receiving second bypass air including a second portion of the propelled air; a core gas path receiving core air including a third portion of the propelled air; a compressor pressurizing the core air; an intercooler facilitating heat removal from the pressurized core air; a combustor in which the pressurized core air is mixed with fuel and ignited to generate a stream of combustion gas; a turbine extracting energy from the combustion gas; and a differential gear train apportioning an input torque received from the turbine between the first fan and the second fan.

Abstract: A fluid cooler for installation in a bypass duct of a turbofan gas turbine engine and associated methods are provided. The fluid cooler includes an inlet duct, a heat exchanger and an outlet duct. The inlet duct includes an inlet protruding into the bypass duct to receive a portion of the bypass air into the inlet duct. The heat exchanger is in fluid communication with the inlet duct. The heat exchanger facilitates heat transfer between a fluid and the portion of bypass air received into the inlet duct. The heat exchanger defines a general flow direction for the portion of bypass air that is different from the main flow direction of bypass air inside the bypass duct. The outlet duct conveys the portion of bypass air from the heat exchanger back to the bypass duct.

Abstract: An oil supply system, has: an oil tank; a closed-loop oil circuit including a component, a pump having a pump inlet and a pump outlet, and a de-aerator; a make-up flow conduit connecting the oil tank to the closed-loop oil circuit; a make-up pump connected to the closed-loop oil circuit; a recirculation conduit in parallel to the make-up pump and bypassing the make-up pump; a fixed orifice fluidly connected to the make-up flow conduit upstream of the closed-loop oil circuit and downstream of the recirculation conduit, the fixed orifice having an area selected to increase a pressure at an outlet of the make-up pump above a pressure at the de-aerator inlet; and a metering orifice connected on the recirculation conduit and having a flow circulating area selected to reduce a flow rate inputted into the closed-loop oil circuit below a flow rate generated by the make-up pump.

Abstract: A cooling system for a compressor casing of an turbofan engine includes a cooling conduit extending between an inlet in fluid communication with a bypass duct of the turbofan engine and an outlet in fluid communication with an outer surface of the compressor casing. A shroud of the compressor is mounted to the compressor casing. A protrusion is disposed about an opening to the inlet and protrudes radially outwardly into the bypass duct A valve in selective fluid communication with the inlet of the cooling conduit is operable to selectively inject a flow of high pressure air from a high pressure air source towards the inlet of the cooling conduit to modulate a flow of bypass air entering the cooling conduit.

Abstract: A system is provided for an aircraft. This aircraft system includes a propulsor, a powerplant and a control system. The powerplant is configured to output mechanical power to drive rotation of the propulsor. The powerplant includes a heat engine and an electric machine. The heat engine is configured to provide a first portion of the mechanical power. The electric machine is configured to provide a second portion of the mechanical power. The control system is configured to operate the powerplant to control a rotational speed of the propulsor by adjusting the second portion of the mechanical power.

Abstract: An aircraft engine is described which has: a shaft; a bearing housing extending around the shaft and defining a bearing cavity; a bearing within the bearing cavity and supporting the shaft; a squeeze film damper having an annulus receiving a damping fluid; and a controller. The controller is configured to: operate the squeeze film damper in a first regime in which a speed of the damping fluid in the annulus is less than a speed of sound in the damping fluid; determine that the shaft is subjected to vibrations having an amplitude greater than an amplitude threshold; and upon determining that the amplitude of the vibrations is greater than the amplitude threshold, operate the squeeze film damper in a second regime in which the speed of the damping fluid in the annulus is greater than the speed of sound in the damping fluid.

Abstract: An engine oil system for an aircraft engine includes an oil tank and an oil tank return. The oil tank forms an internal oil cavity. The oil tank return forms a portion of an oil flow path through the engine oil system. The oil tank return includes a return body forming a return passage extending between and to a passage inlet and a passage outlet. The return body includes an inner diameter wall and an outer diameter wall. The inner diameter wall and the outer diameter wall form a turn of the return passage between the passage inlet and the passage outlet. The return body forms a top side, a corner, and an oblique side at the passage outlet. The corner and the oblique side are disposed at the outer diameter wall. The corner extends between the top side and the oblique side.

Abstract: An engine oil system for an aircraft engine includes an oil tank and an oil tank return. The oil tank forms an internal oil cavity. The oil tank return forms a portion of an oil flow path through the engine oil system. The oil tank return includes a return body forming a return passage extending between and to a passage inlet and a passage outlet. The return body includes an inner diameter wall and an outer diameter wall. The inner diameter wall and the outer diameter wall form a turn of the return passage between the passage inlet and the passage outlet. The return body forms a top side, a corner, and an oblique side at the passage outlet. The corner and the oblique side are disposed at the outer diameter wall. The corner extends between the top side and the oblique side.

Abstract: An aircraft assist system described herein includes an aircraft coupling counterpart attached to a strut of a landing gear of an aircraft, and an assist vehicle. The assist vehicle includes a frame, ground-engaging wheels mounted to the frame, a power source for driving one or more of the ground-engaging wheels, and a vehicle coupling counterpart for engagement with the aircraft coupling counterpart. The aircraft coupling counterpart and the vehicle coupling counterpart define a swivel connection for transferring a propulsive force from the takeoff assist vehicle to the aircraft. The aircraft coupling counterpart is disengageable from the vehicle coupling counterpart by upward movement of the aircraft coupling counterpart relative to the vehicle coupling counterpart.

Abstract: An assembly for an aircraft includes a first shaft, a second shaft and a metering device. The metering device includes a device sidewall, a device first end, a device second end, a device bore, a plurality of channels and a plurality of ports. The device sidewall circumscribes the first shaft at the device first end. The device sidewall is disposed in a second shaft bore of the second shaft at the device second end. The device bore extends axially along the axis through the metering device between the device first end and the device second end. The channels are arranged circumferentially about the axis. Each of the channels projects radially into the device sidewall from the device bore and axially into the device sidewall from the device second end. The ports are arranged circumferentially about the axis. Each of the ports projects radially through the device sidewall.

Abstract: Gas turbine engines include an engine core having compressor, combustor, and turbine sections arranged within an engine casing. A transfer tube is arranged on an exterior of the casing and configured to define a fluid conduit for fluid to be transferred about the engine. A fitting is connected to the transfer tube and a seal assembly is arranged between the fitting and the transfer tube. The seal assembly provides a fluid seal between the fluid conduit and an ambient environment at a connection between the fitting and the transfer tube. The seal assembly includes a primary seal arranged in a gap between the fitting and the transfer tube and a secondary seal arranged in the gap between the primary seal and an ambient environment. The secondary seal is configured to expand into engagement with the fitting and the transfer tube when exposed to temperatures of at least 800° F.

Abstract: An accessory power unit assembly includes a combustor assembly that includes a pulse detonation tube that defines a curved passage between an inlet that receives a compressed airflow and a discharge end. The detonation tube operates as a resonant cavity to sustain a standing pressure wave. Fuel injected into the pulse detonation tube is timed to coincide with movement of a standing wave toward a discharge end. The standing wave provides an increase in pressure of an output exhaust gas driving the turbine without a corresponding increase in load.

Abstract: A de-aerator has: a housing defining an air-oil inlet, an oil outlet, and an air outlet; a rotor within the housing and rotatable about a rotation axis and having: inner and outer rims and blades, passages between the blades and communicating with the air-oil inlet; a conduit extending around the rotation axis and defining a central passage extending axially and communicating with the air outlet, the conduit being radially spaced apart from the outer rim by radial passages communicating with the oil outlet; and openings located inwardly of the blades and communicating with the central passage, wherein one or more of: a cross-sectional area of the central passage taken on a plane transverse to the rotation axis increasing in an axial direction away from the air-oil inlet; and the openings located at least partially radially between an inner face of the conduit and an outer face of the inner rim.

Abstract: A heat management system for a thermal engine includes a source of hydrogen, a first hydrogen conduit flowing a first flow of hydrogen therethrough, the first hydrogen conduit fluidly coupling the source of hydrogen to a fuel system of the thermal engine, and a second hydrogen conduit flowing a second flow of hydrogen therethrough. The second hydrogen conduit is in thermal communication with an engine exhaust conduit to transfer heat from the engine exhaust conduit to the second hydrogen conduit. The second hydrogen conduit is also in thermal communication with the first hydrogen conduit to transfer heat from the second hydrogen conduit to the first hydrogen conduit.

Abstract: An aircraft engine, has: an oil circuit fluidly connecting an oil reservoir to components of the aircraft engine; and a de-aerator fluidly connected to the oil circuit and located downstream of the components relative to a flow of the oil through the oil circuit, the de-aerator feeding the flow of the oil into the reservoir, the de-aerator having: a nozzle extending from a nozzle inlet to a nozzle outlet along a flow passage delimited by a peripheral wall, the peripheral wall extending around a central axis of the nozzle, a flow circulating area of the flow passage decreasing from the nozzle inlet to the nozzle outlet; and a deflector projecting from the peripheral wall of the nozzle at the nozzle outlet, the deflector having a curved section extending between a proximal end at the nozzle outlet and a distal end, the deflector intersected by the central axis of the nozzle.

Abstract: An aircraft assist system described herein includes an aircraft coupling counterpart attached to a strut of a landing gear of an aircraft, and an assist vehicle. The assist vehicle includes a frame, ground-engaging wheels mounted to the frame, a power source for driving one or more of the ground-engaging wheels, and a vehicle coupling counterpart for engagement with the aircraft coupling counterpart. The aircraft coupling counterpart and the vehicle coupling counterpart define a swivel connection for transferring a propulsive force from the takeoff assist vehicle to the aircraft. The aircraft coupling counterpart is disengageable from the vehicle coupling counterpart by upward movement of the aircraft coupling counterpart relative to the vehicle coupling counterpart.

Abstract: An electric motor is provided that includes a housing, a stator, a rotor, and a coolant. The rotor has an output shaft. The rotor is disposed radially inside of the stator and is configured for rotation relative to the stator about a rotational axis. The stator includes an outer radial ring portion and a plurality of stator poles. Each stator pole extends radially inward from the outer radial ring portion to a distal end. The stator includes a plurality of stator cooling circuits. Each stator cooling circuit includes a circumferentially extending outer radial channel located in the outer radial ring portion of the stator and a stator pole channel disposed in each stator pole. The stator pole channel is in fluid communication with the outer radial channel. The stator pole channel extends radially within the respective stator pole. The coolant has magnetic properties disposed in the stator cooling circuits.

Abstract: An assembly is provided for a gas turbine engine. This assembly includes a rotating structure of the gas turbine engine, a stationary structure of the gas turbine engine and a volume formed by and extending between the rotating structure and the stationary structure. The rotating structure is rotatable about an axis. The rotating structure includes a seal land and a fluid nozzle. The stationary structure includes a seal element. The seal element is arranged with the seal land to form a seal assembly that seals an annular gap between the stationary structure and the rotating structure. The volume is adjacent the seal element and the seal land. The fluid nozzle is configured to direct a fluid jet out from the rotating structure and into the volume.

Abstract: A fluid cooler for installation in a bypass duct of a turbofan gas turbine engine and associated methods are provided. The fluid cooler includes an inlet duct, a heat exchanger and an outlet duct. The inlet duct includes an inlet protruding into the bypass duct to receive a portion of the bypass air into the inlet duct. The heat exchanger is in fluid communication with the inlet duct. The heat exchanger facilitates heat transfer between a fluid and the portion of bypass air received into the inlet duct. The heat exchanger defines a general flow direction for the portion of bypass air that is different from the main flow direction of bypass air inside the bypass duct. The outlet duct conveys the portion of bypass air from the heat exchanger back to the bypass duct.

Abstract: An aircraft engine, comprising: a shaft; a bearing housing extending around the shaft and defining a bearing cavity; a bearing in the bearing cavity and rotatably supporting the shaft; a squeeze film damper including: a first seal proximate a first end of the bearing, and a second seal proximate a second end of the bearing; an annulus between the first seal and the second seal, the annulus having a damping fluid in the annulus; and wherein at least the first seal includes a sealing ring received within a groove extending axially from an outer groove wall to an inner groove wall, the inner groove wall located axially between the outer groove wall and the annulus, the sealing ring having an outer ring face facing the outer groove wall and an inner ring face facing the inner groove wall, the inner ring face being movable relative to the outer ring face.