A gas turbine engine for an aircraft including: a compressor stage; a bleed air line diverted from the compressor stage; and a thermoelectric generator system, the thermoelectric generator system including a thermoelectric generator; a vortex tube and a radiator system, the vortex tube including a flow input fluidically connected to the bleed air line to receive an input of compressed gas from the bleed air line, wherein the vortex tube is configured to separate the compressed gas into a hot flow discharged from a first output of the vortex tube, and a cold flow discharged from a second output of the vortex tube; the radiator system including a first heat exchanger and a second heat exchanger disposed on opposing sides of the thermoelectric generator; and at least one of the hot flow is directed to the first heat exchanger, and the cold flow is directed to the second heat exchanger.