Counter rotating aircraft gas turbine engine with high overall pressure ratio compressor

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 second booster being counter rotatable with respect to said high pressure compressor,
• 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 high pressure compressor including between six and eight high pressure stages.

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,
• an outlet guide vane assembly located directly aft of said low pressure turbine,
• 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.

An aircraft gas turbine engine family comprising:

• at least two different engine models having substantially the same fan diameter,
• a first one of the engine models having a one stage high pressure turbine,
• a second one of the engine models having a two stage high pressure turbine,
• each of said models having;
• a high pressure rotor including one of said one and two stage high pressure turbines,
• 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, and
• 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.