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.

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.

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 comprising an inner wall facing the axis X and an outer 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 to allow the hot air flow to escape after heating the inner cavity, the ventilation orifice comprising an upstream edge, the circumferential profile of which is discontinuous in order to generate turbulences, and a downstream edge, the radial profile of which is aerodynamic in order to limit the formation of pressure fluctuations.

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 comprising an inner wall facing the axis X and an outer 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 to allow the hot air flow to escape after heating the inner cavity, the ventilation orifice comprising an upstream edge, the circumferential profile of which is discontinuous in order to generate turbulences, and a downstream edge, the radial profile of which is aerodynamic in order to limit the formation of pressure fluctuations.

A system for lubricating components of a gas turbine engine includes a supply pump to pump lubricant from a reservoir through a supply conduit to lubricated component. An outer body deicing circuit is fluidly coupled in series with the supply conduit. Moreover, a scavenge conduit extends from the lubricated component to the reservoir and a scavenge pump pumps the lubricant from the lubricated component through the scavenge conduit. In addition, a valve is fluidly coupled in series with the supply conduit and the scavenge conduit and a bypass conduit extends from the valve to a location on the supply conduit downstream of the deicing circuit. When the lubricant is supplied to the valve, the valve directs the lubricant flowing through the scavenge conduit into the reservoir. When the supply of lubricant to the valve is halted, the valve directs the lubricant flowing through the scavenge conduit into the bypass conduit.

An airfoil for a fan section of a turbine engine may include a fan blade or an outlet guide vane, and an edge guard attached thereto. The edge guard may include a heating conduit disposed within at least a portion of the edge guard. An anti-icing system for a plurality of fan blades or outlet guide vanes may include a fluid supply pathway configured to supply heating fluid to respective ones of a plurality of heating conduits within the edge guards attached to respective ones of a plurality of fan blades and/or to a plurality of outlet guide vanes. The heating fluid may include bleed air from a core air flowpath. A method of inhibiting icing on an airfoil may include flowing a heating fluid into a heating conduit disposed within an edge guard attached to the airfoil and heating the edge guard with the heating fluid.

An apparatus for aircraft anti-icing includes a nozzle body, at least one nozzle extending from the nozzle body, and at least one vane disposed in at least one of the nozzle(s), the at least one vane configured to impart rotational movement of a hot gas moving through the nozzle(s).

A booster splitter for a gas turbine engine and a method of additively manufacturing the booster splitter are provided. The booster splitter includes an annular inner wall defining a radially outer boundary of a compressor flow path defined through a compressor section of the gas turbine engine, an annular outer wall spaced apart from the annular inner wall along the radial direction, the annular outer wall adjacent to the annular inner wall at a forward end, the forward end defining an inlet to the compressor flow path, an annular bulkhead spanning between the annular inner wall and the annular outer wall substantially along the radial direction, the bulkhead defining an inlet port, and a passageway defined within the annular outer wall, the passageway extending from the inlet port, into the bulkhead, radially outward to the outer wall, and through the annular outer wall towards the inlet defined by the forward end.

A booster splitter for a gas turbine engine and a method of additively manufacturing the booster splitter are provided. The booster splitter includes an annular inner wall defining a radially outer boundary of a compressor flow path defined through a compressor section of the gas turbine engine, an annular outer wall spaced apart from the annular inner wall along the radial direction, the annular outer wall adjacent to the annular inner wall at a forward end, the forward end defining an inlet to the compressor flow path, an annular bulkhead spanning between the annular inner wall and the annular outer wall substantially along the radial direction, the bulkhead defining an inlet port, and a passageway defined within the annular outer wall, the passageway extending from the inlet port, into the bulkhead, radially outward to the outer wall, and through the annular outer wall towards the inlet defined by the forward end.