Gas turbine engine

A gas turbine engine for an aircraft comprising:

• an engine core comprising:
• a turbine, a combustor, and a compressor, 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 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:
• the second turbine comprises at least one ceramic matrix composite component;
• the mass of ceramic matrix composite in the second turbine is in the range of from 1% to 20% of the total mass of the second turbine;
• a cooling flow requirement is defined as the ratio of the mass flow rate of the turbine cooling flow to the mass flow rate of the flow entering the engine core at cruise conditions;
• a turbine entry temperature is defined as the temperature (K) at the inlet to the most axially upstream turbine rotor in the gas turbine engine at a maximum power condition of the gas turbine engine;
• and
• the cooling efficiency ratio, defined as the ratio between the turbine entry temperature and the cooling flow requirement, is in the range of from 8000 K to 20000 K.