Abstract: A direct drive electrically-geared turbofan is provided via a first magnetic gearbox assembly connected to a fan of a turbofan engine; a second magnetic gearbox assembly connected to a spool shaft of the turbofan engine; and a speed controller configured to adjust a rotational speed of the fan based on a rotational speed of the spool shaft by selectively coupling and decoupling the first magnetic gearbox assembly with the second magnetic gearbox assembly. In various aspects, the first or second magnetic gearbox assembly includes a permanent magnet array, while a different one of the first or second magnetic gearbox assemblies includes a rotor winding separated from the permanent magnet array by an air gap; and the speed controller is configured to selectively couple and decouple the first and second magnetic gearbox assemblies with each other via controlling a switch in a winding circuit with the rotor winding.

Abstract: A gas turbomachine combustion chamber includes bridges extending side by side to connect in one piece a radially inner wall and a radially outer wall towards a free end of the radially inner wall. The bridges, inner wall, and outer wall have an additive layer structure.

Abstract: A system may include a first duct including an air inlet end and an air outlet end, and a valve within the first duct and configured to open to allow or close to prevent fan duct air from the air inlet end to pass through to the air outlet end of the first duct. The system may also include a second duct including a first end and a second end. The first end of the second duct may be coupled to a sidewall of the first duct and configured to allow the fan duct air to flow from the first duct to the second duct. The system may also include a resonance chamber coupled to the second end of the second duct and configured to allow the air in the resonance chamber to act as a spring causing the air in the second duct to oscillate at a predefined frequency.

Abstract: An aircraft comprises a machine body. The machine body encloses a turbofan gas turbine engine and a plurality of ancillary systems. The turbofan gas turbine engine comprises, in axial flow sequence, a heat exchanger module, a fan assembly, a compressor module, a turbine module, a combustor module, and an exhaust module. The machine body comprises either one or two fluid inlet apertures. The or each fluid inlet aperture is configured to allow a fluid flow to enter the machine body and to pass through the heat exchanger module. The heat exchanger module is configured to transfer a waste heat load from the gas turbine engine and the ancillary systems to the fluid flow prior to an entry of the fluid flow into the fan module, and thence into the compressor module, the turbine module, the combustor module, and the exhaust module.

Abstract: A gas turbine engine and engine mount structure includes a core engine including a compressor section, a combustor section and a turbine section mounted within a core engine housing. The fan, the compressor section and the turbine section rotate about an axis of rotation. An outer nacelle surrounds the fan, and is spaced from the core engine housing to define a bypass duct. The fan delivers air into the bypass duct and into the core engine housing. The nacelle is formed with camber so as to be curved in a first plane away from the axis of rotation in a first lateral direction. An engine mount structure extends from the nacelle at an angle that is non-parallel and non-perpendicular to the first plane, and has a component in a lateral direction that is opposed to the first lateral direction. An aircraft is also disclosed.

Abstract: A gas turbine engine includes a bypass duct extending about a longitudinal centerline of the gas turbine engine. The bypass duct includes at least one bypass duct wall defining at least a portion of a bypass flow path through the bypass duct. The at least one bypass duct wall includes a scoop extending into the bypass flow path. The gas turbine engine further includes a bleed conduit including an inlet connected to the bypass duct within the scoop of the at least one bypass duct wall and at least one louver mounted to the bleed conduit within the inlet. The at least one louver extends between a leading edge and a trailing edge opposite the leading edge. The leading edge is located within the bypass flow path and the trailing edge is located within the bleed conduit.

Abstract: A gas turbine engine includes a first spool, a second spool, a third spool, and a epicyclic differential. The epicyclic differential includes a sun gear, planet gears supported from a planet carrier, a ring gear, and a locking member, each housed within an accessory gearbox. The planet carrier couples to the second spool. The sun gear couples to the third spool, and the ring gear couples to engine accessories. A method of operating the gas turbine engine includes supplying a first fuel flow rate to a primary combustor associated with the first and third spools and supplying a second fuel flow rate to a secondary combustor associated with the second spool based on a power lever angle within an intermediate power level range.

Abstract: A dual-engine system of an aircraft that includes at least one main engine is provided. The dual-engine system includes a sustainer engine, a booster engine configured to augment a power output of the sustainer engine, and a gearbox coupled to the sustainer engine and the booster engine. An output of the gearbox is configured to selectively combine the power output of the sustainer engine with the booster engine.

Abstract: A gas turbine engine system includes a gas turbine engine, an accessory gearbox, and a flywheel powered barring engine system. The gas turbine engine includes a rotor and a turbine. The accessory gearbox is connected with the rotor. The flywheel powered barring engine system is connected with the accessory gearbox and configured to rotate the rotor after shut down of the gas turbine engine.

Abstract: A disclosed drive assembly for an auxiliary power unit includes a first drive shaft driven by an auxiliary power unit, a second drive shaft driven by a first electric motor, a differential gear system including a ring gear driven by the second drive shaft, a and planet gears supported within a carrier attached to the ring gear. The first drive shaft and an output shaft are coupled to the planet gears and a generator is driven by the output shaft. The electric motor and the auxiliary power unit combine to drive the output shaft.

Abstract: A method of determining an inlet total air pressure includes determining a first parameter indicative of a first inlet total air pressure. The method includes executing a sequence that includes: determining a mass air flow passing through the air inlet based on the first parameter, determining a Mach number of air passing through the air inlet based on the mass air flow, determining a static air pressure at the air inlet, determining an air pressure ratio based on the Mach number, generating a subsequent parameter indicative of the revised inlet total air pressure based on the air pressure ratio and the static air pressure, and substituting the subsequent parameter for the first parameter. The method includes executing at least one additional instance of the sequence with the subsequent parameter, and outputting the subsequent parameter as the inlet total air pressure.

Abstract: A gas turbine engine is provided that includes an engine core and a bypass duct. The engine core includes a compressor section, a combustor section, a turbine section and a core flowpath. The compressor section includes a radial flow compressor rotor. The core flowpath extends through the compressor section, the combustor section and the turbine section from a core inlet to a core exhaust. The bypass duct includes a bypass flowpath that extends outside of the engine core from a bypass inlet to a bypass exhaust. The bypass inlet is disposed along the compressor section and fluidly coupled with core flowpath.

Abstract: A pilot nozzle for a dual fuel turbine engine includes an inner air circuit, a gaseous fuel circuit radially outward from the inner air circuit, a liquid fuel circuit radially outward from the inner air circuit, an outer air circuit radially outward from the liquid fuel circuit and the gaseous fuel circuit, and a shroud radially outward from the outer air circuit. The shroud is configured to stabilize a pilot re-circulation zone downstream from outlets of the inner and outer air circuits and the liquid and gaseous fuel circuits.

Abstract: A gas turbine engine comprises an engine core and an electric fan array. The engine core includes a compressor, a combustor, and a turbine. The compressor compresses and delivers air to the combustor. The combustor mixes fuel with the compressed air received from the compressor and ignites the fuel. The hot, high-pressure products of the combustion reaction in the combustor are directed into the turbine to cause the turbine to rotate about an axis and drive the compressor. The electric fan array may include at least two non-concentric fans having parallel shafts.

Abstract: A gas turbine engine according to an example of the present disclosure may include, among other things, a fan section including a fan having a plurality of fan blades and including an outer housing surrounding the fan blades to establish a bypass duct, a geared architecture, a first spool including a first shaft that interconnects a first compressor and a fan drive turbine, the fan drive turbine driving the fan through the geared architecture. The gas turbine engine is rated to provide an amount of thrust at ground idle, and the gas turbine engine is rated to provide an amount of thrust at maximum takeoff. A thrust ratio is defined as a ratio of the amount of thrust at ground idle divided by the amount of thrust at maximum takeoff. The thrust ratio can be less than or equal to 0.050.

Abstract: A combustion chamber includes a bottom wall having at least one opening, at least one sleeve mounted upstream of the bottom wall and fixed to the bottom wall, a closing ring defining an annular groove with the sleeve and fixed to the sleeve, and at least one air and fuel injection system having an axis, mounted in the opening of the bottom wall. The injection system has an annular flange extending radially with respect to the axis and is mounted in the groove with radial clearance. A baffle is situated downstream of the bottom wall and is fixed to the sleeve and/or to the bottom wall.

Abstract: An aircraft power system includes an auxiliary power unit (APU), a transmission unit, and a generator. The APU is configured to output a first APU rotational power in response to determining a first condition and a second APU rotational power in response to a determining a second condition. The transmission unit is configured to receive the first APU rotational output power from the APU. The transmission unit outputs a transmission rotational power based on the first APU rotational output power and outputs the transmission rotational power based on the second APU rotational output power. The generator receives the transmission rotational power and produces an alternating current (AC) voltage having a target frequency based on the transmission rotational power.

Abstract: The invention provides various designs of an apparatus and method for attemperating a gas stream temperature. The apparatus of the present invention provides a body through which a gas stream passes that permits, as desired, a second gas, such as gas outside of the gas duct or such as ambient air, to be added to the main gas stream to attemperate the temperature of the main gas stream. The body or device may be referred to as a variable eductor having a plurality of openings through which a second gas may pass into the main gas stream. The openings may be opened or closed, and the variable eductor provides control over which openings and the degree to which each opening is opened. In some designs the variable eductor is inserted between two portions of a gas duct. The variable eductor has widespread application, such as downstream of a gas turbine to attemperate the exhaust gas temperature during startup.

Abstract: A bleed plenum for a compressor section associated with a gas turbine engine includes a plenum chamber having a first wall spaced apart from a second wall to define a plenum slot configured to receive a flow of fluid from the compressor section. A slot angle of the plenum slot is defined between a centerline of the plenum slot and a longitudinal axis of the gas turbine engine and varies about a circumference of the bleed plenum. The second wall includes a bullnose having a wedge angle defined between the second wall and a surface associated with the compressor section. The bullnose faces into the flow of fluid through the plenum slot, and the wedge angle of the bullnose varies about the circumference of the bleed plenum.

Abstract: Turbine engine systems and aircraft having such systems are described. The turbine engine systems include a combustor arranged along a core flow path of the turbine engine, a drive shaft having at least a compressor section and a turbine section coupled thereto, a cryogenic fuel tank configured to supply a fuel to the combustor, and an expansion turbine mechanically coupled to the drive shaft, the expansion turbine configured to receive fuel from the cryogenic fuel tank and expand said fuel, wherein expansion of said fuel by the expansion turbine drives rotation of the expansion turbine to provide power input to the drive shaft.