Disclosed is an air intake of an aircraft turbine engine nacelle having a lip, a downstream portion and an internal partition separating the lip and the downstream portion, the lip delimiting an annular recess, the downstream portion having a downstream inner wall and a downstream outer wall, the air intake having an injection channel for injecting a hot air stream into the annular recess, a passage opening formed in the internal partition, an outlet opening formed in the downstream outer wall and a discharge channel for discharging the hot air stream mounted in the downstream portion and having a first end connected to the internal partition, a second end connected to the downstream outer wall and a main body having at least one flexible portion.

A process for the production of an acoustic treatment panel that includes an acoustically resistive layer (24), at least one alveolar structure (26), and a reflective layer (28), as well as at least one channel that is bordered at least in part by a wall (32) that is secant with at least one wall of a cell of the alveolar structure (26), includes elevating the temperature of the wall (32) and the ends of the cells of the alveolar structure (26) and injecting a pressurized gas into the channel (30) so as to expand the wall (32) of the channel so as to flatten it against the ends of the cells of the alveolar structure (26) in such a way as to produce a weld by continuous diffusion between the ends of the walls of the cells of the alveolar structure (26) and the wall (32) of the channel (30).

An inlet for an aircraft nacelle may comprise a nanoreinforced polyimide composite lip skin. The nanomaterials may increase thermal conductivity and decrease microcracking in the lip skin. A lip skin for an inlet with an electric heater may comprise a surface layer, an outer composite skin, an electric heater, an inner composite skin, and a thermal barrier coating. A lip skin for an inlet with a pneumatic deicing system may comprise a surface layer, a composite skin, and a thermal barrier coating.

An engine bleed control system for a gas turbine engine of an aircraft is provided. The engine bleed control system includes an engine bleed tap coupled to a fan-air source or a compressor source of a lower pressure compressor section before a highest pressure compressor section of the gas turbine engine and a turbo-compressor in fluid communication with the engine bleed tap. The engine bleed control system also includes a controller operable to selectively drive the turbo-compressor to boost a bleed air pressure as pressure augmented bleed air and control delivery of the pressure augmented bleed air to an aircraft use.

An apparatus for providing foreign object debris protection and anti-icing capabilities to an air intake of an aircraft engine. The apparatus includes a frame having at least a portion configured to conduct fluids. A tube is positioned such that different portions of the frame are fluidly connected. There is a discharge outlet defined within the frame that is configured to discharge fluids from the frame into the air intake. An inlet is defined within the frame and is fluidly connected to a compressor section of the aircraft engine. In this manner, air from the compressor section of the aircraft engine can flow through the inlet, through the tube, and through the discharge outlet, to return to the compressor section of the aircraft engine.

A device for de-icing a splitter nose of an aviation turbine engine, the device including a splitter nose having an outer annular wall defining the inside of the bypass stream flow channel and an inner annular wall defining an inlet of the core stream flow channel, and an inner shroud mounted at its upstream end on the inner annular wall of the splitter nose and designed to have inlet guide vanes fastened thereto, the splitter nose and the inner shroud defining an annular volume. The device includes an annular deflector positioned inside the annular volume so as to subdivide the annular volume into a first annular cavity and a second annular cavity, the second annular cavity being defined between the annular deflector and the outer annular wall of the splitter nose.

An engine bleed control system for a gas turbine engine of an aircraft is provided. The engine bleed control system includes an engine bleed tap coupled to a fan-air source or a compressor source of a lower pressure compressor section before a highest pressure compressor section of the gas turbine engine and a motorized compressor in fluid communication with the engine bleed tap. The engine bleed control system also includes a controller operable to selectively drive the motorized compressor to boost a bleed air pressure as pressure augmented bleed air and control delivery of the pressure augmented bleed air to an aircraft use.

A die-welding system for a de-icer assembly includes a die, a die base, a high energy source, and a de-icer assembly. The de-icer assembly includes a first welded-material layer and a second welded-material layer. At least one of the die and the die base includes a welded-portion pattern thereon configured to weld the first welded-material layer to the second welded-material layer in the pattern of the welded-portion pattern such that inflatable portions are formed within the welded-portion pattern formed in the de-icer assembly between non-welded sections of the first welded-material layer and the second welded-material layer.

A method of manufacturing a de-icer assembly includes positioning a first welded-material layer and a second welded-material layer between a die and a die base of a die-based welding system, wherein at least one of the die and the die base includes a welded-portion pattern configured to weld the first welded-material layer to the second welded-material layer in the pattern such that inflatable portions are formed within the welded-portion pattern formed in the de-icer assembly between non-welded sections of the first welded-material layer and the second welded-material layer, pressing the first welded-material layer and the second welded-material layer together between the die and die base, and applying high energy to the die-based welding system using a high energy source such that the first welded-material layer and the second welded-material layer are welded together at the areas in the shape of the welded-portion pattern to form a welded de-icer assembly.

Assemblies are provided for an aircraft with a gas turbine engine. One assembly includes a body such as an inlet nose lip of a nacelle for the gas turbine engine. This assembly also includes a thermal management system that includes a duct, a regulator and a flow restrictor. The thermal management system is configured to direct a flow of bleed gas through the duct from the gas turbine engine to the body for substantially preventing ice buildup on the body. The regulator is configured to affect the flow of bleed gas downstream of the regulator. The flow restrictor is configured to selectively restrict the flow of bleed gas through the duct when a temperature of the flow of bleed gas is greater than a threshold temperature.