Rolls-royce Corporation

Paul R Davies

David A Jones

David Reginald Trainer

Stephen Mark Husband

David Peter Scothern

Stephen John Bradbrook

Benjamin J Sellers

Gareth Edward Moore

Luke George

Alexander T Joyce

Andrew J Newman

Gordon Margary

Natalie C Wong

Richard G Mochrie

Douglas M M Herbert

Jonathan A Cherry

Edward C Gower

Douglas Herbert

A thermal management system for an aircraft includes a first thermal bus including one or more first heat sources, a heat sink, a vapour compression system, and one or more second heat sources. The vapour compression system includes a compressor, a condenser, a receiver, a first side of a recuperator, an expansion valve, an evaporator, a second side of the recuperator, and the compressor. A first heat flow (Q) of waste heat energy generated by the first heat sources is transferred via the first heat transfer fluid to the heat sink. A second heat flow (Q) of waste heat energy generated by the second heat source(s) being transferred via the evaporator to a refrigerant. A third heat flow (Q) of heat energy in the refrigerant is transferred via the condenser to the first heat transfer fluid. The controller is configured to ensure that:1.1*<3.0*.

Thermal management system for an aircraft

A thermal management system for an aircraft includes a first gas turbine engine, one or more first electric machines, first thermal bus and heat exchanger. The first thermal bus includes a first heat transfer fluid in a closed loop flow sequence, between the first gas turbine engine, or each first electric machine, and the first heat exchanger. Waste heat energy transfers to the first heat transfer fluid. The first heat exchanger configures to transfer waste heat energy from the first heat transfer fluid to a dissipation medium. During steady-state operation of the first gas turbine engine, the first heat transfer fluid entering the first heat exchanger has a temperature of T(° C.), and a temperature of an inlet air flow entering the first gas turbine engine has a temperature T(° C.) and a ratio B is in a range of between 5.0-18.0.

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 thermal management system for an aircraft comprises a first gas turbine engine, a first thermal bus, a first heat exchanger, and a chiller. The first thermal bus comprises a first heat transfer fluid, with the first heat transfer fluid being in fluid communication, in a closed loop flow sequence, between the first gas turbine engine, the first heat exchanger, and the chiller. Waste heat energy generated by the first gas turbine engine, is transferred to the first heat transfer fluid. The chiller is configured to lower a temperature of the first heat transfer fluid prior to the first heat transfer fluid being circulated through the gas turbine engine. The first heat exchanger is configured to transfer the waste heat energy from the first heat transfer fluid to a dissipation medium.

A gas turbine engine for an aircraft 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 a maximum electrical power P(W), and the gas turbine engine is configured to generate a maximum dry thrust T (N); and a ratio S of: is in a range of between 2.0 and 10.0.

A thermal management system for an aircraft includes a first gas turbine engine, first thermal bus, first heat exchanger, one or more first ancillary systems, vapour compression system, one or more second ancillary systems and second heat exchanger. A waste heat energy generated by a first gas turbine engine, and a first ancillary system, transfers to the first heat transfer fluid. A waste heat energy generated by a second ancillary system transfers to a second heat transfer fluid, and the second heat exchanger transfers the waste heat energy from the second heat transfer fluid to the first heat transfer fluid. The waste heat energy generated by a second ancillary system transfers to the first heat transfer fluid, and the first heat exchanger transfers the waste heat energy to a dissipation medium. The waste heat energy transferred to the second heat transfer fluid ranges from 20 kW to 300 kW.

Multi-engine aircraft power and propulsion systems and methods of restarting an engine of a multi-engine aircraft during fight are provided. One such method comprises: determining a condition to the effect that a flame in the combustion equipment of the second gas turbine engine has been extinguished; responsive to the determination, supplying electrical power from the electrical energy storage system to one or more of the electric machines of the second gas turbine engine and operating said one or more electric machines as motors to limit a reduction in a speed of the one or more spools of the second gas turbine engine following extinguishment of the flame in its combustion equipment; and restarting the second gas turbine engine by relighting the combustion equipment of the second gas turbine engine.

Restarting a 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.

An aircraft gas turbine engine comprises at least a first main engine shaft an electric machine coupled to the main engine shaft, and a lubrication pump configured to provide oil to at least a part of the electric machine. The lubrication pump is controlled by a controller. The controller is configured to operate the lubrication pump in a main gas turbine engine running mode, in which the lubrication pump is driven to provide lubrication to the electric machine at a first flow rate relative to the rotational rate of the main engine shaft, and a fault mode, in which the lubrication pump is driven to provide lubrication to the electric machine at a second, higher flow rate relative to the rotational rate of the main engine shaft.

Gas turbine engine generator oil pump

An aircraft gas turbine engine () comprises a main engine shaft (), a main engine shaft bearing arrangement () configured to rotatably support the main engine shaft () and an electric machine () comprising a rotor () and a stator (). The rotor () is mounted to the main engine shaft () and is rotatably supported by the main engine shaft bearing arrangement (), and the stator () is mounted to static structure () of the gas turbine engine ().

Gas turbine engine electrical generator

An aircraft gas turbine engine includes a main engine shaft, a main engine shaft bearing arrangement for rotatably supporting the main engine shaft, and an electric machine including a rotor and a stator. The rotor is mounted to the main engine shaft and is rotatably supported by a further electric machine bearing arrangement extending between the rotor and the stator, and the stator is mounted to static structure of the gas turbine engine.

A gas turbine engine includes a main gas flow exhaust nozzle having an annular inner surface which, in use, bounds a flow of exhaust gas. The gas turbine engine further includes cooling passages having respective outlets therefrom to provide a flow of cooling air over a surface of the engine or an adjacent airframe component, thereby protecting the cooled surface from the exhaust gas flow. Adjacent cooling passages of the or each pair of the nested cooling passages are separated from each other by a respective dividing wall. The outlets from the nested cooling passages are staggered in the axial direction of the exhaust nozzle such that cooling air flowing out of an inner one of the adjacent cooling passages of the or each pair of the nested cooling passages flows over the dividing wall separating the adjacent passages.

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Bristol, United Kingdom

Paul R Davies