Aircraft engine temperature control and capability

A gas turbine engine for an aircraft comprising:

• an engine core comprising:
• a first turbine a first compressor, and a first core shaft 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, 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:
• part of the flow that enters the engine core bypasses the combustor and is used as turbine cooling flow to cool the 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 to 20000 K.