A gas turbine engine comprises a fan rotor having fan blades received within an outer nacelle, and the outer nacelle having an inner surface. A distance is defined between an axial outer end of the nacelle, and a leading edge of the fan blade. An anti-icing treatment is provided to an inner periphery of the nacelle over at least 75% of the distance along the inner periphery of the nacelle.

An air inflow lip of a nacelle for a turbojet of an aircraft, including a cavity defined by a leading edge of the nacelle and by an annular wall, includes an inner wall, an acoustic treatment device, and a pneumatic de-icing device including a de-icing fluid supply device. The inner wall of the air inflow lip includes acoustic boreholes and the pneumatic de-icing device includes a honeycombed de-icing plate mounted inside the cavity on the inner wall of the air inflow lip. Also included are conduits for the circulation of a de-icing fluid and acoustic wells communicating with the acoustic boreholes.

The turbofan engine can have a core engine, a bypass duct surrounding the core engine, an annular bypass flow path between the bypass duct and the core engine, and a plurality of core links extending across the bypass path and supporting the core engine relative to the bypass duct, and a fluid passage having a heat exchange portion in a given one of the core links, the heat exchange portion being configured for heat exchange with the bypass flow path, an inlet leading into the given core link and to the heat exchange portion, and an outlet extending from the heat exchange portion and out of the given core link.

The turbofan engine can have a core engine, a bypass duct surrounding the core engine, an annular bypass flow path between the bypass duct and the core engine, and a plurality of core links extending across the bypass path and supporting the core engine relative to the bypass duct, and a fluid passage having a heat exchange portion in a given one of the core links, the heat exchange portion being configured for heat exchange with the bypass flow path, an inlet leading into the given core link and to the heat exchange portion, and an outlet extending from the heat exchange portion and out of the given core link.

An aircraft includes a fuselage defining a longitudinal axis between a forward end and an aft end. The aircraft includes an electrical system having an electric storage. The electric storage is positioned within the fuselage.

A system comprises a heat source. The system also comprises a bladder comprising opposing thin-walled sheets and a fluid flow conduit defined between the opposing thin-walled sheets. The fluid flow conduit comprises an inlet and an outlet. The system further comprises a first fluid line coupled to the heat source and the inlet of the bladder. The system additionally comprises a second fluid line coupled to the heat source and the outlet of the bladder. The system also comprises fluid flowable through the first fluid line from the heat source to the inlet, from the inlet through the fluid flow conduit to the outlet, and through the second fluid line from the outlet to the heat source.

A system comprises a heat source. The system also comprises a bladder comprising opposing thin-walled sheets and a fluid flow conduit defined between the opposing thin-walled sheets. The fluid flow conduit comprises an inlet and an outlet. The system further comprises a first fluid line coupled to the heat source and the inlet of the bladder. The system additionally comprises a second fluid line coupled to the heat source and the outlet of the bladder. The system also comprises fluid flowable through the first fluid line from the heat source to the inlet, from the inlet through the fluid flow conduit to the outlet, and through the second fluid line from the outlet to the heat source.

A two-phase type heat transfer device (10) for heat sources operating at a wide temperature range. The heat transfer device (10) includes an evaporator (21) collecting heat from a heat source, a condenser (21) providing heat to a cold sink by a first working fluid passing through liquid and vapor transport lines (25, 27) that connect the evaporator (21) and the condenser (23). The evaporator (21) is arranged inside a saddle (31) configured for avoiding that the temperature of the first working fluid in the evaporator (21) is greater than its critical point. The invention also refers to aircraft ice protection systems using the heat transfer device (10).

A two-phase type heat transfer device (10) for heat sources operating at a wide temperature range. The heat transfer device (10) includes an evaporator (21) collecting heat from a heat source, a condenser (21) providing heat to a cold sink by a first working fluid passing through liquid and vapor transport lines (25, 27) that connect the evaporator (21) and the condenser (23). The evaporator (21) is arranged inside a saddle (31) configured for avoiding that the temperature of the first working fluid in the evaporator (21) is greater than its critical point. The invention also refers to aircraft ice protection systems using the heat transfer device (10).

There is provided a pressure regulating shut-off valve comprising a valve body, at least one piston serving as a regulating piston and/or a shut-off piston, a solenoid valve, and a pressure relief valve; wherein the valve body defines an inlet and an outlet, and comprises at least a portion formed by an additive manufacturing process.