Patent ID: 6966189

Claim:
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.