Patent ID: 8544278

Claim:
A bypass turbojet engine for an aircraft, comprising, around its longitudinal axis (L-L): a nacelle ( 2 ) provided with a nacelle outer cowl ( 3 ) and containing a fan ( 8 ) generating the cold flow ( 9 ) and a central generator ( 10 ) generating the hot flow ( 11 ); an annular cold flow duct ( 12 ) created around said central hot flow generator ( 10 ); a fan outer cowl ( 14 ) delimiting said annular cold flow duct ( 12 ) on said nacelle outer cowl ( 3 ) side; a cold flow outlet orifice ( 6 ) of which the edge ( 7 ), which forms the trailing edge of said nacelle ( 2 ), is determined by said nacelle outer cowl ( 3 ) and by said fan outer cowl ( 14 ) converging toward one another until they meet; and in the vicinity of said cold flow outlet orifice ( 6 ), a plurality of bosses ( 20 ) distributed at the periphery of said fan outer cowl ( 14 ), projecting into said annular cold flow duct ( 12 ) and forming, for said cold flow ( 9 ), a convergent face portion followed by a divergent face portion connected to the edge ( 7 ) of said cold flow outlet orifice ( 6 ), wherein each boss ( 20 ) has a convex face ( 22 ) forming said convergent face portion ( 22 C) and said divergent face portion ( 22 D) and two planar lateral faces ( 20 L), which run longitudinally with respect to said turbojet engine, said convex face ( 22 ) and said lateral faces ( 20 L) giving said boss ( 20 ) an at least approximately rectangular cross section that evolves in a direction parallel to said longitudinal axis (L-L), wherein a peripheral width (l 20 ) of the bosses ( 20 ) is equal to a peripheral width (l 21 ) of a plurality of longitudinal ducts ( 21 ) delimited between the bosses ( 20 ).