Patent ID: 8209953
Filing Date: 2012-07-03
Classification: B64D,F02C,F02K,F04D,Y10T

Abstract:
1. A gas turbine engine system for an aircraft, comprising: a nacelle defined about an axis, wherein said nacelle includes a fan cowl having an inlet lip section and a core cowl disposed at least partially within said fan cowl; at least one compressor and at least one turbine positioned downstream of said at least one compressor; at least one combustor positioned between said at least one compressor and said at least one turbine; a bleed passage having an inlet that receives a bleed airflow and an outlet that discharges said bleed airflow in an upstream direction that is directly opposite of a direction of incoming airflow from said outlet and near a boundary layer of said inlet lip section; and a controller that identifies an operability condition of the aircraft, wherein said controller selectively introduces said bleed airflow near said boundary layer of said inlet lip section of said nacelle in response to said operability condition.