Gas turbine engine
An aircraft gas turbine engine includes a heat exchanger module, and a core engine. The core engine includes an intermediate-pressure compressor, high-pressure compressor, and high and low-pressure turbines. The high-pressure compressor rotationally connects to the high-pressure turbine by a first shaft, and the intermediate-pressure compressor rotationally connects to the low-pressure turbine by a second shaft. The heat exchanger module fluidly communicates with the core engine by an inlet duct. The heat exchanger module includes a central hub and multiple heat transfer elements extending radially from the hub and spaced in a circumferential array, for heat energy transfer from a first fluid within the elements to an inlet airflow passing over a surface of the elements prior to airflow entry into an inlet to the core engine. The gas turbine engine further includes a first electric machine rotationally connected to the first shaft, and positioned downstream of the heat exchanger module.
1. A gas turbine engine for an aircraft, the gas turbine engine comprising:
a core engine comprising, 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 being rotationally connected to the high-pressure turbine by a first shaft, and
the intermediate-pressure compressor being rotationally connected to the low-pressure turbine by a second shaft;
a heat exchanger module comprising:
a central hub, and
a plurality of heat transfer elements extending radially outwardly from the central hub and spaced in a circumferential array, and configured to transfer 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 heat exchanger module being in fluid communication with the core engine by an inlet duct; and
a first electric machine rotationally connected to the first shaft and positioned downstream of the heat exchanger module, the first electric machine having:
a maximum electrical power output of between 180 kW and 500 kW,
an axial length in the range of between 0.20m and 0.45m, and
a maximum diameter in the range of between 0.15m and 0.39m
wherein the first electric machine is positioned axially upstream of the core engine and configured to rotate about a rotational axis of the core engine.
2. The gas turbine engine as claimed in claim 1 , wherein the first electric machine has a mass in the range of between 45 kg and 90 kg.
3. The gas turbine engine as claimed in claim 1 , further comprising:
a second electric machine rotationally connected to the second shaft.
4. The gas turbine engine as claimed in claim 3 , wherein the second electric machine is accommodated within the central hub.
5. The gas turbine engine as claimed in in claim 3 , wherein the second electric machine is positioned axially downstream of the low-pressure turbine.
6. The gas turbine engine as claimed in claim 3 , further comprising:
an electrical energy storage unit configured to store electrical energy generated by at least one of the first electrical machine and the second electrical machine.
7. The gas turbine engine as claimed in claim 6 , wherein the electrical energy storage unit is a battery.
8. The gas turbine engine as claimed in claim 6 , wherein the electrical energy storage unit is a capacitor.
9. The gas turbine engine as claimed in claim 3 , further comprising:
a fan assembly rotationally connected to the second shaft,
wherein the fan assembly, a compressor module comprising the intermediate-pressure compressor and the high-pressure compressor, and a turbine module comprising the high pressure turbine and the low-pressure turbine are in sequential arrangement.
10. The gas turbine engine as claimed in claim 9 , wherein the second electrical machine is rotationally connected to the fan assembly.
11. The gas turbine engine as claimed in claim 9 , wherein:
the fan assembly comprises a plurality of fan blades extending radially from a hub, and
the plurality of fan blades define a fan diameter (D FAN ) that is within the range of 0.3m to 2.0m.
12. The gas turbine engine as claimed in claim 9 , further comprising:
an outer casing enclosing the sequential arrangement of fan assembly, compressor module, and turbine module,
wherein:
an annular bypass duct is defined between the outer casing and the sequential arrangement of compressor module, and turbine module, and
a bypass ratio, defined as a ratio of a mass air flow rate through the bypass duct to a mass air flow rate through the sequential arrangement of compressor module and turbine module is less than 4.0.
13. The gas turbine engine as claimed in claim 9 , wherein:
the fan assembly has two or more fan stages, and
at least one of the fan stages comprises a plurality of fan blades defining a fan diameter D FAN .
14. An aircraft comprising a gas turbine engine as claimed in claim 1 .