Gas turbine engine with improved heat management

A method of operating a gas turbine engine for an aircraft, the method comprising providing the gas turbine engine comprising:

• an engine core comprising a compressor, a combustor, a turbine, and a core shaft connecting the turbine to the compressor;
• a fan comprising a plurality of fan blades and arranged upstream of the engine core;
• turbomachinery bearings;
• a power gearbox adapted to drive the fan at a lower rotation speed than the turbine;
• and
• a heat management system configured to provide lubrication and cooling to the gearbox and turbomachinery bearings, and comprising a pipe assembly including a first lubricant circuit and a second lubricant circuit and adapted to provide a lubricant flow to the gearbox and turbomachinery bearings, at least one first air-lubricant heat exchanger arranged in the first lubricant circuit, and at least one second air-lubricant heat exchanger and at least one fuel-lubricant heat exchanger arranged in the second lubricant circuit, wherein the at least one first air-lubricant heat exchanger and the at least one second air-lubricant heat exchanger are configured to dissipate a first amount of heat to air and the at least one fuel-lubricant heat exchanger is configured to dissipate a second amount of heat to fuel;
• wherein a first proportion of the heat generated by the gearbox and the turbomachinery bearings and dissipated to air is defined as
• (
• first
• ⁢
• amount
• ⁢
• of
• ⁢
• heat
• first
• ⁢
• amount
• ⁢
• of
• ⁢
• heat
• +
• second
• ⁢
• amount
• ⁢
• of
• ⁢
• heat
• )
• 85
• ⁢
• %
• ⁢
• MTO
• at 85% of a core shaft maximum take-off speed;
• and
• wherein the method comprises controlling flow through the at least one first air-lubricant heat exchanger and the at least one fuel-lubricant heat exchanger to provide the first amount of heat and the second amount of heat such that at an environment temperature of ISA +10° C. the first proportion is from 0.40 to 0.60.