A gas turbine engine (10) for an aircraft comprising: an engine core (11) comprising a turbine system comprising one or more turbines (17, 19), a compressor system comprising one or more compressors (14,15), and a core shaft (26) connecting the turbine system to the compressor system, wherein a compressor exit temperature (T30) is defined as an average temperature of airflow at the exit of the highest pressure compressor of the compressor system at cruise conditions and a compressor exit pressure (P30) is defined as an average pressure of airflow at the exit of the highest pressure compressor of the compressor system at cruise conditions, the engine core (11) further comprising an annular splitter (70) at which flow is divided between a core flow (A) that flows through the engine core, and a bypass flow (B) that flows along a bypass duct (22), wherein stagnation streamlines (110) around the circumference of the engine (10), stagnating on a leading edge of the annular splitter (70), form a streamsurface (110) forming a radially outer boundary of a streamtube that contains all of the core flow (A); a fan (23) located upstream of the engine core (11), the fan comprising a plurality of fan blades (64) extending from a hub, each fan blade (64) having a leading edge (64a) and a trailing edge (64b), each fan blade (64) having a radially inner portion (65a) lying within the streamtube that contains the core flow (A), and wherein a fan root entry temperature (T20) is defined as an average temperature of airflow across the leading edge (64a) of the radially inner portion of each fan blade (64) at cruise conditions and wherein a fan root entry pressure (P20) is defined as an average pressure of airflow across the leading edge (64a) of the radially inner portion of each fan blade (64) at cruise conditions; and a nacelle (21) surrounding the engine core (11), the nacelle (21) defining the bypass duct (22) and a bypass exhaust nozzle (18). An overall pressure ratio is defined as the compressor exit pressure (P30) divided by the fan root entry pressure (P20). A bypass nozzle pressure ratio is defined as the nozzle pressure ratio of the bypass exhaust nozzle at cruise conditions. A core temperature rise is defined as the compressor exit temperature (T30) in Kelvin divided by the fan root entry temperature (T20) in Kelvin. A temperature-pressure ratio defined as: the core temperature rise the bypass nozzle pressure ratio is in a range between 1.52 and 1.8, and the overall pressure ratio is in a range between 42.5 and 70. A method of operating the gas turbine engine is also disclosed.