Abstract: A gas turbine, in particular an aircraft engine, including a core flow channel (K), in which a first compressor (20), a second compressor (40) adjacent downstream from the first compressor, a combustion chamber (60) adjacent downstream from the second compressor, a second turbine (50) adjacent downstream from the combustion chamber, which is coupled to the second compressor, and a first turbine (30) adjacent downstream from the second turbine, which is coupled to the first compressor via a first transmission (71), are situated; a quotient (r/R) of an inside diameter (r) of the core flow channel divided by an outside diameter (R) of the core flow channel at an upstream inflow of the first compressor being at most 0.65, in particular at most 0.5 is provided.