An assembly is provided for a structure. This assembly includes a composite skin and a thermal anti-icing system. The composite skin extends between an exterior surface and an interior surface. The thermal anti-icing system includes a susceptor and a waveguide. The susceptor and the waveguide are integrated into the composite skin between the exterior surface and the interior surface. The waveguide is configured to direct microwaves to the susceptor for melting and/or preventing ice accumulation on the exterior surface.

An assembly is provided for an aircraft structure. This aircraft structure assembly includes an acoustic panel and a thermal anti-icing system. The acoustic panel includes an exterior surface. The thermal anti-icing system includes a susceptor and a microwave system. The susceptor is configured with the acoustic panel. The microwave system is configured to direct microwaves to the susceptor for melting and/or preventing ice accumulation on the exterior surface.

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.

There is provided a system and a method for operating a multi-engine rotorcraft. When the rotorcraft is cruising in an asymmetric operating regime (AOR) at least one engine is an active engine and is operated in an active mode to provide motive power to the rotorcraft and at least one second engine is a standby engine and is operated in a standby mode to provide substantially no motive power to the rotorcraft, at least one of a power level of the at least one second engine is increased and at least one variable geometry mechanism of the at least one second engine is moved to shed any ice accumulation on the at least one second engine.

An anti-icing protection system for an aircraft engine nacelle, the nacelle comprising an inner shroud provided with at least one acoustic panel, an air intake lip forming a leading edge of the nacelle, the protection system comprising a heat exchanger device including at least one heat pipe configured to transfer heat emitted by a heat source to the acoustic panel or panels.

The invention relates to an air intake of an aircraft turbojet engine nacelle, extending along an axis X, in which an air flow circulates from upstream to downstream, the air intake extending circumferentially around the axis X and comprising an inner wall, which faces the axis X in order to guide an inner air flow, and an outer wall, which is opposite the inner wall, for guiding an external air flow, the walls being connected by a leading edge and an inner partition so as to delimit an annular cavity. The air intake comprises means for injecting at least one hot air flow into the inner cavity and at least one ventilation orifice formed in the outer wall in order to allow the hot air flow to escape after heating the internal cavity, the air intake comprising at least one disruption member of the external air flow, positioned upstream of the ventilation orifice, which extends outwardly from the outer wall.

A turbofan engine is provided. The turbofan engine includes a fan comprising a plurality of fan blades; a turbomachine operably coupled to the fan for driving the fan, the turbomachine comprising a compressor section, a combustion section, and a turbine section in serial flow order and together defining a core air flowpath; a nacelle surrounding and at least partially enclosing the fan; an inlet pre-swirl feature located upstream of the plurality of fan blades, the inlet pre-swirl feature attached to or integrated into the nacelle; and a means for reducing ice buildup or ice formation on the inlet pre-swirl feature, the means in communication with the inlet pre-swirl feature.

According to the present teachings, an aircraft engine configured to controllably shed ice during icing conditions is presented. The engine has a member operating at a first angular velocity having a surface divided into a first ice accumulating surface configured to collect ice, and a first shadow surface configured to resist the collection of ice. A first flange is disposed between the first ice accumulating surface and the first shadow surface.

An assembly is provided for an aircraft propulsion system. The assembly includes a nacelle inlet structure with an internal cavity. The assembly also includes a nozzle configured to direct fluid into the internal cavity through a plurality of ports that include one or more first ports and at least one second port. The nozzle includes a trunk conduit, a first branch conduit and a second branch conduit. The first branch conduit and the second branch conduit are fluidly coupled in parallel to the trunk conduit. The first branch conduit includes the first port(s). The second branch conduit includes the second port.

An air inlet section for an enclosure for an aircraft engine is provided that includes an inner barrel panel, an outer barrel panel, a lipskin and a forward bulkhead. The lipskin extends between an inner barrel end and an outer barrel end. The inner barrel end is disposed proximate the forward end of the inner barrel panel and the outer barrel end is disposed proximate the forward end of the outer barrel panel. The forward bulkhead has a panel that extends between an outer radial end and an inner radial end. The inner barrel panel, the outer barrel panel, and the lipskin define an interior annular region, and the forward bulkhead defines a sub-portion of interior annular region. The outer radial end of the forward bulkhead panel is disposed forward of the inner radial end of the forward bulkhead panel.