An air-breathing turbojet engine (101) for a hypersonic vehicle is shown. The engine comprises a pump for pumping a cryogenic fuel, an inlet (102) configured to compress inlet air by one or more shocks, a cooler (103) to cool the compressed inlet air using the cryogenic fuel, and a turbo-compressor (104) to compress the air further. A combustor (105) receives compressed cooled air from the turbo-compressor and a first portion of the cryogenic fuel for combustion. A first turbine (106) expands and is driven by combustion products, and a second turbine (107) expands and is driven by a second portion of the cryogenic fuel. The first turbine and the second turbine drive the turbo-compressor via a shaft. An afterburner (109) receives combustion products from the first turbine and the second portion of the cryogenic fuel from the second turbine for combustion therein.

An aircraft power plant comprising novel air management features for high-power fuel cell applications, said features combine supercharging and turbocharging elements with air and hydrogen gas pathways, utilize novel airflow concepts and provide for much stronger integration of various fuel cell drive components.

A rotorcraft having an aerodynamic device arranged below a rotor, which rotor participates at least in providing lift for the rotorcraft, the rotor being mounted to rotate about an axis of rotation, the aerodynamic device having a fairing provided with at least one air inlet for enabling a stream of cool air to flow from a region that is situated outside the rotorcraft to another region that is situated inside, the air inlet being delimited by an outer peripheral portion of the fairing. In accordance with the invention, the aerodynamic device has a perforated plate provided with at least one perforation, the perforation being suitable for allowing the stream of cool air to pass through it, the perforated plate having at least one main portion shaped to match an outer shape of the outer peripheral portion of said fairing.

An aircraft propulsor assembly includes a fan having a nacelle and plural fan blades radially disposed within the nacelle. The fan blades are configured to be rotated by torque generated by a turbine engine of an aircraft to generate thrust for propelling the aircraft. The assembly also includes an electric motor including a stator in the nacelle of the fan and a rotor in tips of two or more of the fan blades. The electric motor is configured to generate torque that also rotates the fan blades to generate thrust for propelling the aircraft. The assembly also includes a controller configured to reduce or prevent an increase in an operating temperature of the turbine engine of the aircraft by automatically supplanting at least some of the torque generated by the turbine engine with the torque generated by the electric motor.

A cabin air compressor (CAC) motor cooling flow enhancement cowl includes a cowl body contoured to extend from a CAC motor cooling duct discharge of a ram air outlet header. The cowl body includes a backward facing step to form an air dam in a ram air flow path to divert ram system air away from the CAC motor cooling duct discharge. The CAC motor cooling flow enhancement cowl also includes a coupling interface formed at an upstream end of the cowl body to engage with a transition edge of the CAC motor cooling duct discharge.

A gas turbine engine includes a fan assembly, a compressor assembly, a combustion chamber, a turbine assembly, a bypass duct conveying rearward a bypass airstream driven by the fan assembly when the gas turbine engine is in use, a fairing extending across at least a portion of the bypass duct downstream of the fan assembly, and a heat exchanger having an inlet fluidly connected to the compressor assembly and an outlet fluidly connected to a pneumatic actuator of the gas turbine engine. The fairing has a leading edge and a trailing edge. The heat exchanger is disposed adjacent the trailing edge of the fairing.

Heat-exchange and noise-reduction panel for a propulsion assembly, in particular for an aircraft, the panel comprising: a perforated plate comprising a plurality of through-openings; a cellular structure comprising longitudinally oriented structural walls covered by said perforated plate and comprising, between said walls, cavities that define Helmholtz resonators, said through-openings forming necks of said resonators; and means for the circulation of fluid, for example oil, at said perforated plate, wherein said fluid circulation means comprise channels that are formed at least in part in thickened ends of said walls on the same side as said perforated plate, and/or at least in part in regions of the perforated plate situated in the longitudinal extension of said thickened ends.

A gas turbine engine coupled to an aircraft includes an engine core arranged axially along an axis, a bypass duct arranged circumferentially around the engine core to define a bypass channel, and a heat exchanger system. The bypass channel is arranged to conduct bypass air around the engine core to provide thrust for the gas turbine engine. The heat exchanger system is configured to provide cooling for the engine core.

An assembly for an aircraft having a propeller, including an engine assembly having an engine shaft configured for driving engagement with the propeller. The engine assembly includes first and second heat exchangers configured for circulation of at least one of a liquid coolant and a lubricant therethrough. A wheel well is configured for receiving a retracted landing gear. A first cooling duct receives the first heat exchanger and has a first outlet downstream of the first heat exchanger, and a second cooling duct receives the second heat exchanger and has a second outlet downstream of the second heat exchanger. The outlets are in direct fluid communication with an environment of the aircraft, and laterally spaced from each other. The wheel well is located between the outlets. A method of cooling an engine assembly is also discussed.

An aircraft including an aft-mounted boundary layer energisation system is shown. The system comprises a nacelle arranged around a tailcone of the aircraft which thereby defines a duct, the duct having, in axial flow series, an intake, a heat exchanger, and a nozzle, and no turbomachinery therein, whereby the system energises a boundary layer of the aircraft by means of Meredith effect.