Patent ID: 8096105

Claim:
A turbofan turbine engine for an aircraft, comprising: a hollow nacelle having a longitudinal axis and comprising, at the front, an air inlet and, at the rear, an air outlet; a fan placed axially in said nacelle opposite said air inlet and capable of generating the cold flow of said turbine engine; a generator placed axially in said nacelle, behind said fan, said generator being capable of generating the axial hot flow of said turbine engine surrounded by said cold flow and being enclosed in an engine cowl; and a fan channel inner cowl coaxially surrounding said hot flow generator so as: to delimit with the nacelle a channel of annular section for said cold flow, the channel terminates in said air outlet of the nacelle; to delimit with said engine cowl an intermediate chamber of annular section; and to converge via a rear portion of the fan channel inner cowl with a rear portion of said engine cowl so that respective rear edges of the rear portions form an edge of an outlet orifice of said hot flow at a rear portion of said intermediate chamber, wherein: provided in the rear portion of said intermediate chamber are communication means placed about said longitudinal axis and configured to place said intermediate chamber in communication with the outside, in the vicinity of a boundary between said cold flow and said hot flow; a plurality of hatches are provided that are arranged in said rear portion of the engine cowl, while being distributed on the periphery of the rear portion of the engine cowl; said hatches are opened only when the speed of said turbine engine is greater than a threshold corresponding to at least the cruising speed of the aircraft; and in the open position, said hatches draw off, from said hot flow, individual jets of hot air flowing into said intermediate chamber before leaving the intermediate chamber through said communication means while being distributed about said longitudinal axis.