A GAS TURBINE ENGINE

There is disclosed a gas turbine engine (10) for an aircraft, the gas turbine engine comprising: a compressor stage; a bleed air line (50) diverted from the compressor stage; and a thermoelectric generator system (100, 200, 300). The thermoelectric generator system comprises a thermoelectric generator (102); a vortex tube (104) and a radiator system (106). The vortex tube comprises a flow input (112) 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 (114) of the vortex tube, and a cold flow discharged from a second output (116) of the vortex tube. The radiator system comprises a first heat exchanger (120) and a second heat exchanger (118) disposed on opposing sides of the thermoelectric generator. 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.