Patent ID: 7993099

Claim:
A helicopter gas turbine engine comprising: a compressor which receives external air through an air inlet channel; a combustion chamber; a first turbine arranged downstream from the combustion chamber which receives combustion gases emitted by the combustion chamber so as to drive the compressor through a first shaft; a second turbine arranged downstream from the first turbine which is connected by to a second shaft to a gear train so as to supply mechanical power to an output shaft, the first shaft and the second shaft being coaxial, and the second shaft and the output shaft are not coaxial; and a nozzle, wherein the nozzle includes a part forming a diffuser connected downstream from the second 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, and wherein said ejector extends in the downstream direction beyond the downstream end of the diffuser and includes a wall formed at least partially by a sound attenuator which attenuates sound frequencies generated by at least one of the rotation of at least one of the first or second turbines or by the combustion chamber, and a downstream end of said ejector is bent away from an axis of the engine.