A gas turbine engine for an aircraft comprises an engine core. The engine comprises a fan located upstream of the engine core. The engine comprises a nacelle surrounding the fan and the engine core and defining a bypass duct, where the 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. The engine comprises a plurality of actuators and a fuel supply system arranged to supply fuel for combustion in the combustor, and to supply fuel to fueldraulically drive at least one actuator. The fuel comprises at least 25% SAF by volume, and the fuel supply system is arranged such that a peak differential pressure of the fuel across the at least one fueldraulic actuator during cruise conditions is at least 2400 kPa.