Gas turbine engine

A gas turbine engine for an aircraft comprising:

• an engine core comprising a turbine, a compressor, a fan located upstream of the compressor and comprising a plurality of fan blades, and a core shaft connecting the turbine to the compressor;
• and
• a gearbox which receives an input from the core shaft and outputs drive, via a drive shaft, to the fan so as to drive the fan at a lower rotational speed than the turbine, the drive shaft and the core shaft forming a shaft system, wherein the shaft system provides:
• a first portion which extends forward from a first thrust bearing to the fan, the first thrust bearing supporting the shaft system and being located between the turbine and the gearbox, and
• a second portion extending rearward from the first thrust bearing to the turbine, such that in the event of a shaft break within the second portion of the shaft system, said shaft break dividing the shaft system into a front portion axially located by the first thrust bearing and a rear portion no longer axially located by the first thrust bearing, the rear portion is free to move axially rearwardly under a gas load;
• wherein the engine further comprises a shaft break detector, configured to detect the shaft break in the shaft system,
• wherein the shaft break detector includes a phonic wheel, located at the first portion of the shaft system, and a microwave sensor, configured to sense a rotational speed of the shaft system at the second portion of the shaft system axially spaced from the first portion, and
• wherein the turbine has a plurality of stages, the microwave sensor being provided at a first stage of the plurality of stages that is closest to the compressor and the core shaft being connected to the turbine at a final stage of the plurality of stages.