A gas turbine engine for an aircraft comprising: an engine core comprising: a turbine (17, 19), a combustor, and a compressor (14, 15), the turbine comprising a first turbine and a second turbine and the compressor comprising a first compressor and a second compressor; a first core shaft connecting the first turbine to the first compressor; a second core shaft connecting the second turbine to the second compressor, the second turbine, second compressor, and second core shaft being arranged to rotate at a higher rotational speed than the first core shaft, the gas turbine engine further comprising: a bypass duct radially outside the engine core; a fan comprising a plurality of fan blades; and 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, wherein: part of the flow that enters the engine core bypasses the combustor and is used as turbine cooling flow to cool the turbine; the fan diameter is in the range of from 225cm to 400cm; at cruise conditions, the cooling to bypass flow efficiency ratio defined as the ratio of the mass flow rate of the turbine cooling flow to the mass flow rate of the bypass flow at engine cruise conditions is not greater than 0.02; and the turbine entry temperature defined as the temperature at the inlet to the most axially upstream turbine rotor at a maximum power condition of the gas turbine engine, is at least 1800K; characterised in that the second turbine comprises first and second rotor blade and stator vane rows, wherein the most axially upstream row of stator vanes and rotor blades of the second turbine are metallic, whilst the second most axially upstream row of stator vanes and rotor blades of the second turbine are manufactured using Ceramic Matrix Composites (CMCs). an engine core comprising: a turbine (17, 19), a combustor, and a compressor (14, 15), the turbine comprising a first turbine and a second turbine and the compressor comprising a first compressor and a second compressor; a first core shaft connecting the first turbine to the first compressor; a second core shaft connecting the second turbine to the second compressor, the second turbine, second compressor, and second core shaft being arranged to rotate at a higher rotational speed than the first core shaft, the gas turbine engine further comprising: a turbine (17, 19), a combustor, and a compressor (14, 15), the turbine comprising a first turbine and a second turbine and the compressor comprising a first compressor and a second compressor; a first core shaft connecting the first turbine to the first compressor; a second core shaft connecting the second turbine to the second compressor, the second turbine, second compressor, and second core shaft being arranged to rotate at a higher rotational speed than the first core shaft, the gas turbine engine further comprising: a bypass duct radially outside the engine core; a fan comprising a plurality of fan blades; and 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, wherein: part of the flow that enters the engine core bypasses the combustor and is used as turbine cooling flow to cool the turbine; the fan diameter is in the range of from 225cm to 400cm; at cruise conditions, the cooling to bypass flow efficiency ratio defined as the ratio of the mass flow rate of the turbine cooling flow to the mass flow rate of the bypass flow at engine cruise conditions is not greater than 0.02; and the turbine entry temperature defined as the temperature at the inlet to the most axially upstream turbine rotor at a maximum power condition of the gas turbine engine, is at least 1800K; characterised in that the second turbine comprises first and second rotor blade and stator vane rows, wherein the most axially upstream row of stator vanes and rotor blades of the second turbine are metallic, whilst the second most axially upstream row of stator vanes and rotor blades of the second turbine are manufactured using Ceramic Matrix Composites (CMCs). part of the flow that enters the engine core bypasses the combustor and is used as turbine cooling flow to cool the turbine; the fan diameter is in the range of from 225cm to 400cm; at cruise conditions, the cooling to bypass flow efficiency ratio defined as the ratio of the mass flow rate of the turbine cooling flow to the mass flow rate of the bypass flow at engine cruise conditions is not greater than 0.02; and the turbine entry temperature defined as the temperature at the inlet to the most axially upstream turbine rotor at a maximum power condition of the gas turbine engine, is at least 1800K; characterised in that the second turbine comprises first and second rotor blade and stator vane rows, wherein the most axially upstream row of stator vanes and rotor blades of the second turbine are metallic, whilst the second most axially upstream row of stator vanes and rotor blades of the second turbine are manufactured using Ceramic Matrix Composites (CMCs).