System for sealing the secondary flow at the inlet to a nozzle of a turbomachine having a post-combustion chamber

1. An aviation turbomachine comprising, downstream from the turbine, a post-combustion chamber extended by at least one nozzle, said chamber being defined radially by a thermal protection lining disposed inside a casing, said casing and said lining together defining an annular channel in which, in operation, there flows a secondary flow, an annular diaphragm secured to said casing being disposed at the downstream end of said channel, said nozzle comprising a plurality of flaps hinged to the upstream end of said casing, each flap being fitted on its inside face with a thermal protection plate co-operating with said flap to define a passage for being fed with cooling air delivered by said diaphragm,

wherein the feed of cooling air to said passages is provided by an annular duct defined on the outside by a first flexible annular gasket retained, in operation, pressed in sliding contact against the downstream inside face of the casing and against the upstream inside faces of the flaps under urging from the pressure of the secondary flow, and defined on the inside by a second flexible annular gasket whose upstream end is fixed to the radially inner zone of the diaphragm, and whose downstream end is pressed in sliding contact against the upstream inside face of the protection plates.

2. A turbomachine according to claim 1 , further including a third flexible annular gasket for providing sealing between the diaphragm and the thermal protection lining, said third gasket being held upstream against the diaphragm with its downstream end bearing slidably against said protection lining.

3. A turbomachine according to claim 1 , wherein each annular gasket is constituted by a plurality of sectors, each comprising two superposed plates connected together and offset in the circumferential direction so that the edges of two adjacent sectors overlap, each plate presenting, downstream, a plurality of axial slots closed by another plate.

4. A turbomachine according to claim 3 , wherein the slots extend over at least half the axial extent of said gaskets.

5. A turbomachine according to claim 3 , wherein the plates of the sectors are connected together by welding or by brazing.

6. A turbomachine according to claim 1 , wherein the diaphragm is constituted by a channel section ring whose flanges extend upstream and whose web includes orifices, the radially outer flange being fixed to the casing by means leaving an annular gap between said flange and said casing, the upstream end of the first gasket being received with clearance in said gap.

7. A turbomachine according to claim 6 , wherein the upstream end of the second gasket is held clamped between a support plate and the radially inside face of the radially inner flange by means of rivets fixing said support plate to said flange.

8. A turbomachine according to claim 7 , further including a third flexible annular gasket for providing sealing between the diaphragm and the thermal protection lining, said third gasket being held upstream against the diaphragm with its downstream end bearing slidably against said protection lining, and wherein the upstream end of the third gasket is fixed to the radially inside face of the support plate by said rivets.

9. A turbomachine according to claim 7 , wherein said rivets have heads bearing slidably against the outside face of the protective lining.

10. A turbomachine according to claim 1 , wherein each thermal protection plate is fixed to the associated flap by means of a single fastener device, said flap and said plate being prevented from moving relative to each other in rotation about said fastener device by an axial slideway and rail system, said protection plate presenting at its upstream end and on its radially inside face a surface that is convex in the axial direction, providing sealing by contact with the downstream end of the second gasket over the entire annular operating range of the nozzle.