A gas turbine engine for an aircraft comprising: an engine core comprising a first turbine , a first compressor , a first core shaft connecting the first turbine to the first compressor, a second turbine , a second compressor , and a second core shaft connecting the second turbine to the second compressor, with the second turbine, second compressor, and second core shaft being arranged to rotate at a higher rotational speed than the first core shaft and the second compressor positioned axially downstream of the first compressor; a fan located upstream of the engine core, the fan comprising a plurality of fan blades; a gearbox that receives an input from the first core shaft and outputs drive to the fan so as to drive the fan at a lower rotational speed than the first core shaft, and a front mount and a rear mount , the front and rear mounts being configured to connect the gas turbine engine to the aircraft, wherein: the front mount is coupled to a casing of the engine core, the casing surrounding the compressor and separating a core airflow (A) and a bypass airflow (B), and the front mount is located forward of the centre of gravity of the gas turbine engine; the front mount is provided in an axial region that overlaps the first compressor ; the front mount is positioned such that a line (L1) passing through the front mount and the centre of gravity (CG) of the gas turbine engine subtends a first angle (α) of 5 degrees or less relative to a plane (P) perpendicular to a longitudinal axis of the gas turbine engine; and the fan has a diameter in the range of from 250 cm to 390 cm. an engine core comprising a first turbine , a first compressor , a first core shaft connecting the first turbine to the first compressor, a second turbine , a second compressor , and a second core shaft connecting the second turbine to the second compressor, with the second turbine, second compressor, and second core shaft being arranged to rotate at a higher rotational speed than the first core shaft and the second compressor positioned axially downstream of the first compressor; a fan located upstream of the engine core, the fan comprising a plurality of fan blades; a gearbox that receives an input from the first core shaft and outputs drive to the fan so as to drive the fan at a lower rotational speed than the first core shaft, and a front mount and a rear mount , the front and rear mounts being configured to connect the gas turbine engine to the aircraft, wherein: the front mount is coupled to a casing of the engine core, the casing surrounding the compressor and separating a core airflow (A) and a bypass airflow (B), and the front mount is located forward of the centre of gravity of the gas turbine engine; the front mount is provided in an axial region that overlaps the first compressor ; the front mount is positioned such that a line (L1) passing through the front mount and the centre of gravity (CG) of the gas turbine engine subtends a first angle (α) of 5 degrees or less relative to a plane (P) perpendicular to a longitudinal axis of the gas turbine engine; and the fan has a diameter in the range of from 250 cm to 390 cm. the front mount is coupled to a casing of the engine core, the casing surrounding the compressor and separating a core airflow (A) and a bypass airflow (B), and the front mount is located forward of the centre of gravity of the gas turbine engine; the front mount is provided in an axial region that overlaps the first compressor ; the front mount is positioned such that a line (L1) passing through the front mount and the centre of gravity (CG) of the gas turbine engine subtends a first angle (α) of 5 degrees or less relative to a plane (P) perpendicular to a longitudinal axis of the gas turbine engine; and the fan has a diameter in the range of from 250 cm to 390 cm. <EOS>