CPC F02C 7/32 (2013.01) [B64D 27/10 (2013.01); B64D 33/08 (2013.01); F05D 2220/323 (2013.01); F05D 2260/606 (2013.01)] | 10 Claims |
1. A turbine engine for an aircraft, comprising:
a supply device of compressed air configured to supply compressed air to the aircraft;
a cooling system of the compressed air supplied to the aircraft, comprising a scoop configured to collect cooling air in a flow duct of a secondary flow; and
a management system of a diameter of a turbine casing, configured to be supplied with the cooling air collected from the scoop of the cooling system, and configured to control radial clearances between the turbine casing and a plurality of turbine rotor vanes tips,
wherein the scoop is located on a tubular arm configured for the passage of a plurality of auxiliaries of an intermediate casing, the tubular arm extending radially between an inner radial end fixed to an inner shell of the intermediate casing and an outer radial end fixed to an outer shell of the intermediate casing, the inner and outer shells of the intermediate casing defining a portion of the flow duct of the secondary flow, the scoop being located at a junction between the tubular arm and an inner annular wall defining an inner face of the flow duct,
wherein the scoop is connected by a bypass to a first supply channel of the management system and a second supply channel of the cooling system, the bypass being located in the tubular arm, wherein the first supply channel passes into the tubular arm and through the inner shell and the second supply channel passes into the tubular arm and through the outer shell;
wherein the scoop includes an opening in direct contact with the secondary flow, the first supply channel and the second supply channel are separate channels that only intersect at the junction at the opening, such that the first and the second supply channels are in direct contact with the secondary flow.
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