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    Turbofan gas turbine engine
    1.
    发明授权
    Turbofan gas turbine engine 有权

    公开(公告)号：US11668235B2

    公开(公告)日：2023-06-06

    申请号：US17484096

    申请日：2021-09-24

    Applicant: ROLLS-ROYCE plc

    Inventor： Natalie C Wong ,  Thomas S Binnington ,  David A Jones ,  Daniel Blacker

    IPC: F02C7/04 ,  F02K3/06 ,  F02C7/14 ,  F02C7/10

    CPC classification number: F02C7/04 ,  F02C7/10 ,  F02C7/14 ,  F02K3/06 ,  F05D2220/32 ,  F05D2260/213

    Abstract: A turbofan gas turbine engine comprises, in axial flow sequence, a heat exchanger module, an inlet duct, a fan assembly, a compressor module, and a turbine module. The fan assembly comprises a plurality of fan blades defining a fan diameter D, and the heat exchanger module comprises a plurality of heat transfer elements for transfer of heat from a first fluid contained within the heat transfer elements to an airflow passing over a surface of the heat transfer elements prior to entry of the airflow into the fan assembly.
    In use, the first fluid has a maximum temperature of 80° C., and the heat exchanger module transfers at least 300 kW of heat energy from the first fluid to the airflow.

    Gas turbine engine
    2.
    发明授权
    Gas turbine engine 有权

    公开(公告)号：US12092030B2

    公开(公告)日：2024-09-17

    申请号：US17940055

    申请日：2022-09-08

    Applicant: ROLLS-ROYCE plc

    Inventor： Natalie C Wong ,  Jonathan A Cherry ,  Paul R Davies ,  David A Jones ,  Andrew J Newman ,  Benjamin J Sellers ,  Stephen J Bradbrook

    IPC: F02C7/12 ,  F02C6/20 ,  F02C7/32

    CPC classification number: F02C7/12 ,  F02C6/20 ,  F02C7/32 ,  F05D2220/323 ,  F05D2220/76 ,  F05D2260/213

    Abstract: 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.

    Thermal management system for an aircraft
    3.
    发明授权
    Thermal management system for an aircraft 有权

    公开(公告)号：US11952945B2

    公开(公告)日：2024-04-09

    申请号：US18233518

    申请日：2023-08-14

    Applicant: ROLLS-ROYCE plc

    Inventor： Paul R Davies ,  Alexander T Joyce ,  David A Jones

    IPC: F02C7/16 ,  F02C6/18 ,  F25B27/02

    CPC classification number: F02C7/16 ,  F02C6/18 ,  F25B27/02 ,  F05D2220/323 ,  F05D2260/213

    Abstract: 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.

    Fuel management system
    4.
    发明授权
    Fuel management system 有权

    公开(公告)号：US11933225B2

    公开(公告)日：2024-03-19

    申请号：US18166285

    申请日：2023-02-08

    Applicant: ROLLS-ROYCE plc

    Inventor： Richard G Mochrie ,  David A Jones ,  Gordon Illing

    IPC: F02C7/224

    CPC classification number: F02C7/224 ,  F05D2220/32 ,  F05D2240/35 ,  F05D2260/213 ,  F05D2270/07

    Abstract: A fuel management system includes: a fuel tank; a fuel supply line that supply fuel from the fuel tank to a combustor of the gas turbine engine; a reheat fuel supply line that supply fuel from fuel tank to a reheat of the gas turbine engine, the reheat fuel supply line extending from a reheat branching point on the fuel supply line to the reheat; a fuel supply pump disposed along fuel supply line upstream of the reheat branching point; a reheat pump disposed along reheat fuel supply line, the reheat pump that pressurise fuel to a reheat delivery pressure for delivery to the reheat; and a reheat recirculation line that recirculate fuel from the reheat fuel supply line to a location upstream of fuel supply pump, the reheat recirculation line extending from a reheat recirculation branching point on the reheat fuel supply line downstream of the reheat pump.

    Gas turbine engine
    6.
    发明授权
    Gas turbine engine 有权

    公开(公告)号：US11994067B2

    公开(公告)日：2024-05-28

    申请号：US17940055

    申请日：2022-09-08

    Applicant: ROLLS-ROYCE plc

    Inventor： Natalie C Wong ,  Jonathan A Cherry ,  Paul R Davies ,  David A Jones ,  Andrew J Newman ,  Benjamin J Sellers ,  Stephen J Bradbrook

    IPC: F02C7/12 ,  F02C6/20 ,  F02C7/32

    CPC classification number: F02C7/12 ,  F02C6/20 ,  F02C7/32 ,  F05D2220/323 ,  F05D2220/76 ,  F05D2260/213

    Abstract: 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.

    Thermal management system for an aircraft
    9.
    发明授权
    Thermal management system for an aircraft 有权

    公开(公告)号：US12116933B2

    公开(公告)日：2024-10-15

    申请号：US18233521

    申请日：2023-08-14

    Applicant: ROLLS-ROYCE PLC

    Inventor： Richard G Mochrie ,  David A Jones

    IPC: F02C7/14 ,  F02C7/16 ,  F02C7/224

    CPC classification number: F02C7/14 ,  F02C7/16 ,  F02C7/224 ,  F05D2220/323 ,  F05D2260/213 ,  F05D2270/303

    Abstract: A thermal management system for an aircraft comprises a first gas turbine engine, one or more first electric machines rotatably coupled to the first gas turbine engine, a first thermal bus, and a first heat exchanger module. 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 or each first electric machine, and the first heat exchanger. Waste heat energy generated by at least one of the first gas turbine engine, and the or each first electric machine, is transferred to the first heat transfer fluid. The first heat exchanger module comprises a first flow path and a second flow path. The first flow path is configured to direct a flow of the first heat transfer fluid either to a first heat dissipation portion in which a first proportion QA of the waste heat energy from the first heat transfer fluid is transferred to a first dissipation medium, or additionally to a second heat dissipation portion in which a second proportion QB of the waste heat energy from the first heat transfer fluid is transferred to a second dissipation medium, in dependence on a temperature of the first heat transfer fluid entering the first heat exchanger module, a temperature of the first heat dissipation medium, and a temperature of the second heat dissipation medium. The second flow path is configured to direct the flow of the first heat transfer fluid to a second heat dissipation portion in which a second proportion QB of the waste heat energy from the first heat transfer fluid is transferred to a second dissipation medium, or additionally to the first heat dissipation portion in which a first proportion QA of the waste heat energy from the first heat transfer fluid is transferred to a first dissipation medium, in dependence on a temperature of the first heat transfer fluid entering the first heat exchanger module, a temperature of the first heat dissipation medium, and a temperature of the second heat dissipation medium.

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