Patent ID: 6622475
Filing Date: 2003-09-23
Classification: F01D,F02C,F04D,Y02T

Abstract:
A bypass gas turbine engine for an aircraft having an intermediate structural casing (9) positioned between a primary flow path (2) bounded by inner and outer walls and a bypass flow path (3), and axially configured between a low pressure compressor (5) and a high pressure compressor (7), said engine further having a bleed device arranged to deflect a portion of a gas flow from the low pressure compressor (5) toward the bypass flow path (3), said bleed device comprising:an annular cavity (16) defining a manifold situated upstream said intermediate casing (9) and opening along the outer wall of the primary flow path (2), said manifold permanently communicating with said primary flow path (2); a plurality of conduits (18) substantially axially aligned with a central axis of said engine and positioned along the intermediate casing (9), said plurality of conduits connecting to said manifold (16); a plurality of tubes (20) configured around the high pressure compressor (7) and connecting said conduits (18) to the bypass flow path (3); and at least one flow regulating valve (22) disposed along the interior of each of said plurality of tubes and cooperating therewith to regulate the gas flow disposed therein.