A gas turbine engine for an aircraft includes a compressor, a combustion chamber, and a turbine having at least one stator, and at least one rotor. Each stator and rotor is formed by a plurality of blades, a fluid channel is formed between two consecutive blades, and each blade has two opposing surfaces. The compressor is in fluid communication with a first group of stator channels, and the combustion chamber is in fluid communication with a second group of stator channels, such that heat exchange can be performed through two opposing surfaces of at least one stator blade. The outer and the inner walls define a duct for the passage of the heated fluid through the rotor blades, and the outer wall is also arranged for directing the compressed air towards the combustion chamber.

A bowed rotor prevention system for a gas turbine engine of an aircraft is provided. The bowed rotor prevention system includes a gear train and a bowed rotor prevention motor. The gear train is coupled through an engine accessory to an engine accessory gearbox that is further coupled to a starting spool of the engine. The bowed rotor prevention motor is operable to drive rotation of the starting spool of the gas turbine engine through the gear train upon engine shutdown.

An engine assembly for an aircraft, including an internal combustion engine having a liquid coolant system in fluid communication with a heat exchanger, an exhaust duct in fluid communication with air passages of the heat exchanger, a fan in fluid communication with the exhaust duct for driving a cooling air flow through the air passages of the heat exchanger and into the exhaust duct, and an intermediate duct in fluid communication with an exhaust of the engine and having an outlet positioned within the exhaust duct downstream of the fan and upstream of the outlet of the exhaust duct. The outlet of the intermediate duct is spaced inwardly from a peripheral wall of the exhaust duct. The engine assembly may be configured as an auxiliary power unit. A method of discharging air and exhaust gases in an auxiliary power unit having an internal combustion engine is also discussed.

A system for driving an environmental control system (ECS) of a vehicle with ground-based electrical power may include a turbine engine on board the vehicle, coupled to the ECS to pneumatically drive the ECS, and an electric machine, on board the aircraft, mechanically coupled to the turbine engine, to drive the turbine engine. A control system such that the electric machine provides motoring assistance to the turbine engine, and that the motoring assistance is limited to match the available current from the ground-based electrical power.

Integrated arrangements of electrical machines and power electronics converters are described. One such arrangement comprises: an electrical machine comprising one or more windings; a power electronics converter arranged to supply current to or receive current from the one or more windings of the electrical machine; a magnetocaloric effect (MCE) material in thermal contact with the power electronics converter; and a heat sink for removing heat from the MCE material. The MCE material is arranged in proximity to the one or more windings of the electrical machine whereby, in use, stray magnetic flux from the windings of the electrical machine passes through the MCE material and activates the MCE material. The repeated application and removal of the stray flux during normal operation of the electrical machine creates cycles of magnetic refrigeration, which removes heat from the power electronics converter.

An airplane has main propulsion engines and a first fuel supply for the main propulsion engines. The airplane further has an auxiliary propulsion engine and a second fuel supply for the auxiliary propulsion engine, this second fuel supply being separate from the first fuel supply. The auxiliary propulsion engine can be switched on independently from the main propulsion engines. Such airplane has increased safety, since it will be possible to maintain flight, particularly when at high altitude, even if all main propulsion engines have failed.

A gas turbine engine for an aircraft comprising an engine core and including a bypass channel which radially surrounds the engine core at least in part is described. A core shaft is operatively connected to an engine accessory gearbox, which is arranged between the engine core and the bypass channel, by means of a radial shaft of a drive train. An electric machine is provided which is designed to start the gas turbine engine during motor operation and to generate electrical energy during alternator operation. The electric machine is arranged coaxially with the core shaft and connected thereto for conjoint rotation. Alternatively, the electric machine can be arranged radially outside the bypass channel and can be operatively connected to the core shaft by means of the radial shaft, wherein a rotor of the electric machine is arranged coaxially with the radial shaft and connected thereto for conjoint rotation.

One example of a mount for a rotorcraft comprises a structural support member, a bracket, and an elastomer. The bracket is configured to attach to a component of the rotorcraft. The component of the rotorcraft produces vibrations at a first frequency. The structural support member configured to transfer a weight of the component of a rotorcraft to an airframe of the rotorcraft. A rotor system of the aircraft vibrates the airframe of the rotorcraft at a second frequency. The elastomer is located between a structural support member and a bracket. The elastomer is configured to attenuate noise caused by the vibrations at the first frequency by isolating the vibrations at the first frequency from reaching the airframe of the rotorcraft while the airframe vibrates at the second frequency.

A method of cooling an aircraft power plant having a combustion engine is disclosed. The method comprises in a first operating mode, inducing a cooling air flow through a heat exchanger in an air conduit via a flow inducing device fluidly connected to the air conduit, the heat exchanger connected in heat exchange relationship with the power plant of the aircraft. The method comprises, in a second operating mode, bypassing the cooling air flow from the flow inducing device via a selectively closable air outlet of the air conduit downstream of the heat exchanger. A cooling system for an aircraft power plant is also disclosed.

A system for supplying electrical energy to at least one electrical load of an aircraft, this system including a main source of electrical energy connected to at least one electrical load through an electrical network, the system including a control device connected to the main source of electrical energy and configured to deliver a variable voltage level on the electrical network based on a predetermined voltage variation law depending on at least one parameter characterizing the altitude of the aircraft and according to the Paschen law with a predefined margin.