Rolls-Royce Corporation

Andrew J Newman

Stephen John Bradbrook

Paul R Davies

Benjamin J Sellers

Gordon Margary

Natalie C Wong

David A Jones

Jonathan A Cherry

An aircraft gas turbine engine includes a heat exchanger module, and a core engine including an intermediate-pressure compressor, a high-pressure compressor, a high pressure turbine, and a low-pressure turbine. The high-pressure compressor is connected to the high-pressure turbine by a first shaft, and the intermediate-pressure compressor is connected to the low-pressure turbine by a second shaft. The heat exchanger module includes a central hub and heat transfer elements extending radially from the central hub and spaced in a circumferential array, for transferring heat energy from a fluid within the heat transfer elements to an inlet airflow passing over the heat transfer elements prior to entry of the airflow into an inlet to the core engine. The gas turbine engine further includes a first electric machine connected to the first shaft and positioned downstream of the heat exchanger module, and a second electric machines connected to the second shaft.

Gas turbine engine

A gas turbine engine for an aircraft includes, in axial flow sequence, a compressor module, a combustor module, and a turbine module, with a first electric machine being rotationally connected to the turbine module. The first electrical machine is configured to generate a total electrical power P(W), and the gas turbine engine is configured to generate a total shaft power P(W); and a ratio R of: is in a range of between 0.005 and 0.020.

Gas turbine engine with electric machines

A cooling system for an aircraft comprises a gas turbine engine, an ancillary apparatus, and a heat exchanger. The gas turbine engine comprises, in axial flow sequence, a compressor module, a combustor module, and a turbine module, with a first electric machine being rotationally connected to the turbine module. The first electrical machine is configured to generate an electrical power P(W). The heat exchanger is configured to transfer a total waste heat energy Q (W) generated by the gas turbine engine and the ancillary apparatus, to an airflow passing through the heat exchanger, and a ratio S of: is in a range of between 0.50 and 5.00.

A blisk for use in the turbine section of a gas turbine engine comprises a disk around the circumference of which is disposed a plurality of blades in an annular array. The blades are preferably cast separately and then joined, for example by welding, so that the roots form a continuous ring. The remaining volume of the rotor disk is formed of consolidated metal powder by a hot isostatic process. Extending a short distance above each blade root is a shank carrying a platform and the blade airfoil section. The dimensions of the platforms in the circumferential direction are such that they co-operate to form a plenum chamber encircling the periphery of the rotor disk. The blades are cast with internal cooling passages access to which is gained from the plenum through orifices in the sides of the blade shanks. Thus, in use, cooling air may be passed across at least one face of the blisk into the plenum from where it enters the blade cooling passages.

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Bristol, United Kingdom

Andrew J Newman