Ventilation system for a convergent divergent exhaust nozzle

1. A bypass turbojet comprising an afterburn chamber of axis X delimited by an annular wall situated radially inside an annular casing, said annular wall and said casing defining an annular passage through which circulates a stream of cooling air, a convergent divergent axisymmetric nozzle arranged downstream of said afterburn chamber and comprising a circle of convergent flaps hinged at the downstream end of said casing and a circle of divergent flaps hinged at the downstream end of said convergent flaps, each circle of flaps comprising alternately a plurality of controlled flaps, and a plurality of follower flaps, a circle of cold flaps arranged radially outside said nozzle and hinged at their upstream end to a conical shell linked to the downstream part of said casing, means for creating a film of cooling air on the internal faces of said convergent flaps and means of cooling said divergent flaps, wherein the means of cooling said divergent flaps comprise:

an annular plenum chamber delimited downstream by said conical shell and fed with cooling air through drillings made in a boundary wall between said plenum chamber and the downstream end of said annular passage,

a plurality of distribution cells surrounding the plenum chamber and linked to the latter, said cells being delimited downstream by said conical shell and being arranged around the X axis in the planes of symmetry of the follower flaps, and

telescopic pipelines each linking a cell to the follower divergent flap situated in the same plane of symmetry as said cell.

2. The turbojet as claimed in claim 1 , wherein the follower divergent flaps are partitioned and cooled by the air delivered by the telescopic pipelines while the controlled divergent flaps have a single skin.

3. The turbojet as claimed in claim 1 , wherein the convergent flaps are of the single skin type.

4. The turbojet as claimed in claim 1 , wherein the conical wall comprises openings between the cells to allow the circulation of a nacelle air in the space surrounding the convergent divergent nozzle.

5. The turbojet as claimed in claim 1 , wherein the stream of cooling air circulating through the annular duct is divided into two streams by means of a stationary ring integral with the boundary wall, the radially inner stream being injected upstream of the convergent flaps via a slot and the radially outer stream being injected into the plenum chamber through the drillings of the boundary wall.

6. The turbojet as claimed in claim 5 , wherein the means for cooling the divergent flaps furthermore comprise means for adjusting the flow rate of cooling air for said flaps.

7. The turbojet as claimed in claim 6 , wherein the flow rate adjustment means comprise a ring mounted movably in a slide integral with the boundary wall, said ring and said slide each comprising a plurality of flow rate adjustment holes capable of being matched up with the drillings of the boundary wall by displacement of said ring.

8. The turbojet as claimed in claim 7 , wherein the ring is mounted movably in rotation about the X axis and is driven in rotation by a rack and pinion system by means of an actuator driving said pinion.

9. The turbojet as claimed in claim 7 , wherein the ring is mounted movably in translation parallel to the X axis and is displaced by a plurality of synchronized jacks.