公开(公告)号：US20170298822A1

公开(公告)日：2017-10-19

申请号：US15378238

申请日：2016-12-14

Applicant: Airbus Operations, S.L.

Inventor： Alexandre Garde La Casa

IPC: F02C7/08 ,  B64D41/00 ,  F01D5/02 ,  F01D9/04 ,  F02C9/18 ,  F02C3/04 ,  B64D27/10 ,  F01D15/12 ,  F01D17/10 ,  F01D15/10

CPC classification number: F02C7/08 ,  B64D27/10 ,  B64D41/00 ,  B64D2041/002 ,  F01D1/12 ,  F01D5/02 ,  F01D9/041 ,  F01D15/10 ,  F01D15/12 ,  F01D17/105 ,  F02C3/04 ,  F02C3/145 ,  F02C7/10 ,  F02C9/18 ,  F05D2220/323 ,  F05D2220/50 ,  F05D2240/35 ,  F05D2240/60 ,  Y02T50/671

Abstract: A gas turbine engine for an aircraft includes a compressor, a combustion chamber, and a turbine having at least one stator, and at least one rotor. Each stator and rotor is formed by a plurality of blades, a fluid channel is formed between two consecutive blades, and each blade has two opposing surfaces. The compressor is in fluid communication with a first group of stator channels, and the combustion chamber is in fluid communication with a second group of stator channels, such that heat exchange can be performed through two opposing surfaces of at least one stator blade. The outer and the inner walls define a duct for the passage of the heated fluid through the rotor blades, and the outer wall is also arranged for directing the compressed air towards the combustion chamber.