Abstract: The soundproofed exhaust duct for a turbine engine comprises a perforated inner shell defining a flow passage, and a solid outer shell, the inner and outer shells defining between them a space that is closed at the upstream and downstream ends of said space. A core extends between the inner and outer shells at a distance therefrom, and contains at least one sound energy dissipating layer made up of hollow beads that are held against one another. A frame with upstream and downstream portions connected together by longitudinal members supports the core and partitions it into a plurality of boxes that are filled with hollow beads that are held between perforated textures, the frame being secured to the outer shell and/or to the inner shell via at least one of its upstream and downstream portions.

Abstract: The soundproofed exhaust duct for a turbine engine comprises a perforated inner shell defining a flow passage, and a solid outer shell, the inner and outer shells defining between them a space that is closed at the upstream and downstream ends of said space. A core extends between the inner and outer shells at a distance therefrom, and contains at least one sound energy dissipating layer made up of hollow beads that are held against one another. A frame with upstream and downstream portions connected together by longitudinal members supports the core and partitions it into a plurality of boxes that are filled with hollow beads that are held between perforated textures, the frame being secured to the outer shell and/or to the inner shell via at least one of its upstream and downstream portions.

Abstract: A helicopter gas turbine engine having a combustion chamber, at least one turbine arranged downstream from the combustion chamber to receive combustion gases emitted by the latter and a nozzle is disclosed. The nozzle has one part forming a diffuser, connected downstream from the turbine and an ejector which has an upstream part surrounding the downstream end of the diffuser and defining therewith an outlet passage for an engine compartment cooling air secondary flow, the ejector running in the downstream direction beyond the downstream end of the diffuser. The ejector has a wall formed at least partially by a sound attenuator calculated to attenuate sound frequencies generated by the rotation of the or each turbine and/or by the combustion chamber.

Abstract: In a helicopter gas turbine engine, lips (30, 32) defining a radial air intake (34) are connected along an opening (12a) of an engine compressor air supply passage (12). The air intake lips are at least partially formed by a sound attenuator (40, 42) calculated to attenuate sound frequencies produced by the compressor rotation.

Abstract: In a helicopter gas turbine engine, lips (30, 32) defining a radial air intake (34) are connected along an opening (12a) of an engine compressor air supply passage (12). The air intake lips are at least partially formed by a sound attenuator (40, 42) calculated to attenuate sound frequencies produced by the compressor rotation.

Abstract: The invention relates to a helicopter gas turbine engine having a combustion chamber, at least one turbine (20) arranged downstream from the combustion chamber to receive combustion gases emitted by the latter and a nozzle. The nozzle has one part forming a diffuser (26), connected downstream from the turbine and an ejector (30) which has an upstream part surrounding the downstream end of the diffuser and defining therewith an outlet passage (32) for an engine compartment cooling air secondary flow, the ejector running in the downstream direction beyond the downstream end of the diffuser. The ejector (30) has a wall formed at least partially by a sound attenuator (34) calculated to attenuate sound frequencies generated by the rotation of the or each turbine and/or by the combustion chamber.