Patent ID: 9188008
Filing Date: 2015-11-17
Classification: F01D,Y02T

Abstract:
1. A gas turbine for aeronautic engines comprising a first and at least a second rotating bladed sector comprising respectively a first and a second turbine disk arranged coaxially to an axis of the turbine and respectively carrying a coupled first and second moving-blade crown, the first and the second turbine disk defining with associated blades first and second passages through which a cooling air mass for said turbine disks can pass, a cooling air passageway conveying cooling air to the first and second passages with the cooling air exiting the first passages and flowing towards the second passages, first and second stator bodies interposed between said first and second moving blade crowns, a first seal interposed between the first moving blade crown and said first stator bodies and a second seal interposed between said second stator bodies and said second blade crowns, and a thermally insulating chamber insulating a hot gas flows from the cooling air passageway, said thermal insulating chamber comprising a mechanical barrier stably connected to said stator bodies and defining with the stator bodies an annular duct to convey a bled flow of hot gas exiting through said first seal towards said second seal.