Gas turbine engine

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

• an engine core comprising a turbine, a combustor, a compressor, and a core shaft connecting the turbine to the compressor;
• a fan located upstream of the engine core and arranged to be driven by the core shaft, the fan comprising a plurality of fan blades;
• a nacelle surrounding the fan and the engine core and defining a bypass duct located radially outside of the engine core, where a bypass ratio, defined as the ratio of the mass flow rate of the flow through the bypass duct to the mass flow rate of the flow through the core at cruise conditions, is at least 4;
• a plurality of actuators comprising a variable compressor vane actuator;
• and
• a fuel supply system;
• and wherein the method comprises:
• supplying, using the fuel supply system, fuel for combustion in the combustor;
• and
• supplying, using the fuel supply system, fuel to fueldraulically drive the variable compressor vane actuator of the plurality of actuators,
• wherein the fuel is thermally stable at temperatures exceeding 280° C.