Gas turbine engine with intermediate case bearing arrangement

A gas turbine engine for an aircraft, the gas turbine engine comprising:

• a rotatable shaft defining a rotational axis, the rotatable shaft extending between a rearward end and a forward end of the gas turbine engine;
• a compressor drum surrounding, and coupled to, the shaft so as to define an annular gap therebetween;
• an intermediate case arranged axially rearward of the compressor drum and comprising a bearing support element extending into the annular gap;
• a rearward bearing mounted between the intermediate case and the shaft;
• a forward bearing that is forward of the rearward bearing, mounted between the bearing support element and the shaft, and proximate a forward end of the compressor drum such that the forward bearing is less than 50% of an axial length of the compressor drum from the forward end of the compressor drum;
• a fan assembly that is forward of the compressor drum, the fan assembly comprising a fan that has fan blades and is located forward of the compressor drum;
• and
• a gearbox arranged to receive an input from the rotatable shaft and output drive to the fan so as to drive the fan at a lower rotational speed than the rotatable shaft;
• wherein the rotatable shaft is a hollow shaft defining an elongate cavity, and a fan thrust shaft is located within the cavity of the rotatable shaft and is coaxial with the rotatable shaft, the fan thrust shaft being connected to the fan of the gas turbine engine.