A gas turbine engine 10 for an aircraft comprises an engine core 11 comprising a turbine 19, a compressor 14, and a core shaft 26 connecting the turbine 19 to the compressor, the turbine 19 being the lowest pressure turbine of the engine 10, and comprising turbine blades 19a-19d, and the compressor 14 being the lowest pressure compressor of the engine; a fan 23 located upstream of the engine core 11; and a gearbox 30 that receives an input from the core shaft 26 and outputs drive to the fan. The engine core further comprises three bearings arranged to support the core shaft 26, the three bearings comprising two rearward bearings 26b, 26c located downstream of the leading edge of the lowest pressure turbine blade 19c, 19d of the turbine 19 at the root of the blade, and/or, when the turbine 19 comprises four sets 19a-d of turbine blades, downstream of the trailing edge of a turbine blade of the third set 19c of turbine blades from the front of the turbine 19, at the root of the blade.


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    Title :

    Shaft bearing arrangement


    Publication date :

    2020-02-05


    Type of media :

    Patent


    Type of material :

    Electronic Resource


    Language :

    English


    Classification :

    IPC:    B64D Ausrüstung für Flugzeuge , EQUIPMENT FOR FITTING IN OR TO AIRCRAFT



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