A gas turbine engine for an aircraft includes an engine core including a turbine, a compressor, a core shaft, and a core exhaust nozzle, the core exhaust nozzle having a core exhaust nozzle pressure ratio calculated using total pressure at the core nozzle exit; a fan including a plurality of fan blades; and a nacelle surrounding the fan and the engine core and defining a bypass duct, the bypass duct including a bypass exhaust nozzle, the bypass exhaust nozzle having a bypass exhaust nozzle pressure ratio calculated using total pressure at the bypass nozzle exit;
wherein a bypass to core ratio of: bypass ⁢ ⁢ exhaust ⁢ ⁢ nozzle ⁢ ⁢ pressure ⁢ ⁢ ratio core ⁢ ⁢ exhaust ⁢ ⁢ nozzle ⁢ ⁢ pressure ⁢ ⁢ ratio is configured to be in the range from 1.1 to 2.0 under aircraft cruise conditions.