AN IMPROVED GAS TURBINE ENGINE

A gas turbine engine for an aircraft comprises, in axial flow sequence, a heat exchanger module, and a core engine. The core engine comprises, in axial flow sequence, an intermediate-pressure compressor, a high-pressure compressor, a high pressure turbine, and a low-pressure turbine. The high-pressure compressor is rotationally connected to the high-pressure turbine by a first shaft, and the intermediate-pressure compressor is rotationally connected to the low-pressure turbine by a second shaft. The heat exchanger module is in fluid communication with the core engine by an inlet duct. The heat exchanger module comprises a central hub and a plurality of heat transfer elements extending radially outwardly from the central hub and spaced in a circumferential array, for transfer of heat energy from a first fluid contained within the heat transfer elements to an inlet airflow passing over a surface of the heat transfer elements prior to entry of the airflow into an inlet to the core engine. The gas turbine engine further comprises a first electric machine. The first electric machine is rotationally connected to the first shaft, the first electric machine is positioned downstream of the heat exchanger module, the first electric machine has a maximum electrical power output of between 180kW and 500kW, the first electric machine has an axial length in the range of between 0.20m and 0.45m, and the first electric machine has a maximum diameter in the range of between 0.15m and 0.39m.