Patent ID: 6732502
Filing Date: 2004-05-11
Classification: F01D,F02K,F04D,F05D,Y02T

Abstract:
An aircraft gas turbine engine comprising:a high pressure rotor including a high pressure turbine, a low pressure turbine having counter rotatable low pressure inner and outer rotors located aft of said high pressure rotor, said low pressure inner and outer rotors including low pressure inner and outer shafts which are at least in part rotatably disposed co-axially with and radially inwardly of said high pressure rotor, said low pressure turbine including a rotatable annular outer drum rotor drivingly connected to a first fan blade row and a first booster by said low pressure inner shaft, said low pressure turbine further including a rotatable annular inner drum rotor drivingly connected to a second fan blade row and a second booster by said low pressure outer shaft, said first and second boosters axially located between said first and second fan blade rows, a bypass duct radially bounded by a fan casing and an annular radially inner bypass duct wall surrounding said first and second boosters, a radially outer portion of said second fan blade row is radially disposed within said bypass duct, a fan inlet hub to tip radius ratio in a range between 0.20 and 0.35, a high pressure compressor of said high pressure rotor drivenly connected to said high pressure turbine by a high pressure shaft, said high pressure compressor being operable to produce an overall pressure ratio in a range of about 40-65, a bypass ratio in a range of 5-15, an operational fan pressure ratio in a range of 1.4-2.5, and a sum of operational fan tip speeds of said first and second fan blade rows in a range of 1000 to 2500 feet per second, and said second booster being counter rotatable with respect to said high pressure compressor.