公开(公告)号：US12209536B2

公开(公告)日：2025-01-28

申请号：US18609514

申请日：2024-03-19

Applicant: ROLLS-ROYCE plc

Inventor： Nicholas Grech ,  Matthew J. Willshee ,  Jacopo Tacconi

IPC: F02C7/224 ,  B64D37/30 ,  B64D37/34 ,  F02C3/22

Abstract: A propulsive aircraft gas turbine engine comprises a turbine disposed in a gas turbine engine core flow and a recuperator heat exchanger disposed downstream of the turbine in gas turbine engine core flow, the recuperator heat exchanger being configured to transfer heat from the gas turbine engine core flow to gas turbine engine fuel. The recuperator heat exchanger is configured to accommodate a portion of the gas turbine engine core flow therethrough, the remainder being bypassed around the recuperator heat exchanger.

公开(公告)号：US12221924B2

公开(公告)日：2025-02-11

申请号：US18379728

申请日：2023-10-13

Applicant: ROLLS-ROYCE PLC

Inventor： Nicholas Grech ,  Chloe J Palmer ,  Jacopo Tacconi

IPC: F02C6/04 ,  F02C7/224 ,  F02C9/40 ,  H01M8/04111

Abstract: A combined gas turbine engine and hydrogen fuel cell system includes a hydrogen fuelled gas turbine engine, a cryogenic liquid hydrogen fuel tank, a first fuel offtake configured and arranged to divert a portion of hydrogen fuel from a main fuel conduit, a burner configured and arranged to burn the portion of hydrogen fuel diverted from the main fuel conduit, a heat exchanger configured and arranged to transfer heat from exhaust gasses produced by the burner to hydrogen fuel in the main fuel conduit, a second fuel offtake arranged to divert a portion of hydrogen fuel from the main fuel conduit downstream of the heat exchanger, and a hydrogen fuel cell configured and arranged to produce electric power using hydrogen fuel diverted from the second fuel offtake.

公开(公告)号：US12163464B2

公开(公告)日：2024-12-10

申请号：US17749908

申请日：2022-05-20

Applicant: ROLLS-ROYCE plc

Inventor： Richard G Stretton ,  Michael C Willmot ,  Nicholas Grech

IPC: F02C3/06 ,  F01D17/16 ,  F02K1/00 ,  F02K3/06

Abstract: A gas turbine engine (10) for an aircraft comprises an engine core (11) comprising a turbine (19), a compressor (14), a core shaft (26), and a core exhaust nozzle (20), the core exhaust nozzle (20) having a core exhaust nozzle pressure ratio calculated using total pressure at the core nozzle exit (56); a fan (23) comprising a plurality of fan blades; and a nacelle (21) surrounding the fan (23) and the engine core (11) and defining a bypass duct (22), the bypass duct (22) comprising a bypass exhaust nozzle (18), the bypass exhaust nozzle (18) having a bypass exhaust nozzle pressure ratio calculated using total pressure at the bypass nozzle exit;
wherein a bypass to core ratio of: bypass ⁢ exhaust ⁢ nozzle ⁢ pressure ⁢ ratio core ⁢ exhaust ⁢ nozzle ⁢ pressure ⁢ ratio is configured to be in the range from 1.1 to 1.4 under aircraft cruise conditions.

公开(公告)号：US12270341B2

公开(公告)日：2025-04-08

申请号：US18379717

申请日：2023-10-13

Applicant: ROLLS-ROYCE PLC

Inventor： Chloe J Palmer ,  Jacopo Tacconi ,  Nicholas Grech

IPC: F02C7/224 ,  F02C3/22 ,  F02C7/14

Abstract: A gas turbine engine fuel system includes a fuel offtake configured to divert a portion of hydrogen fuel from a main fuel conduit, a burner configured to burn the portion of hydrogen fuel diverted from the main fuel conduit, a heat exchanger configured to transfer heat from exhaust gasses produced by the burner to hydrogen fuel in the main fuel conduit and a power recovery device configured to extract power from exhaust gasses of the burner downstream of the heat-exchanger.

公开(公告)号：US20250035037A1

公开(公告)日：2025-01-30

申请号：US18913201

申请日：2024-10-11

Applicant: ROLLS-ROYCE PLC

Inventor： Richard G. Stretton ,  Michael C. Willmot ,  Nicholas Grech

IPC: F02C3/06 ,  F01D17/16 ,  F02K1/00 ,  F02K3/06

Abstract: A gas turbine engine (10) for an aircraft comprises an engine core (11) comprising a turbine (19), a compressor (14), a core shaft (26), and a core exhaust nozzle (20), the core exhaust nozzle (20) having a core exhaust nozzle pressure ratio calculated using total pressure at the core nozzle exit (56); a fan (23) comprising a plurality of fan blades; and a nacelle (21) surrounding the fan (23) and the engine core (11) and defining a bypass duct (22), the bypass duct (22) comprising a bypass exhaust nozzle (18), the bypass exhaust nozzle (18) having a bypass exhaust nozzle pressure ratio calculated using total pressure at the bypass nozzle exit;
wherein a bypass to core ratio of: bypass ⁢ exhaust ⁢ nozzle ⁢ pressure ⁢ ratio core ⁢ exhaust ⁢ nozzle ⁢ pressure ⁢ ratio is configured to be in the range from 1.1 to 2.0 under aircraft cruise conditions.

公开(公告)号：US12092024B1

公开(公告)日：2024-09-17

申请号：US18601182

申请日：2024-03-11

Applicant: ROLLS-ROYCE plc

Inventor： Nicholas Grech ,  Jacopo Tacconi

IPC: F02C6/18 ,  F02C9/28

CPC classification number: F02C6/18 ,  F02C9/28

Abstract: A gas turbine engine includes a core engine having a compressor, a combustor and a turbine in sequential air flow series. The engine further includes a fuel offtake configured and arranged to divert a portion of hydrogen fuel from a main fuel conduit, a burner configured and arranged to burn the portion of hydrogen fuel diverted from the main fuel conduit and a heat exchanger configured and arranged to transfer heat from exhaust gasses produced by the burner to hydrogen fuel in the main fuel conduit. At least first and second compressor bleed offtakes are at different pressure stages of the compressor, each being configured to bleed a portion of air from the compressor. At least first and second compressor bleed offtake valves are configured to control flow through the first and second bleed offtakes respectively. The burner is configured to receive bleed air from the compressor bleed offtakes.

公开(公告)号：US11339713B2

公开(公告)日：2022-05-24

申请号：US16398742

申请日：2019-04-30

Applicant: ROLLS-ROYCE plc

Inventor： Richard G Stretton ,  Michael C Willmot ,  Nicholas Grech

IPC: F02K3/06 ,  F02K1/00 ,  F02C3/06 ,  F01D17/16

Abstract: A gas turbine engine for an aircraft includes an engine core including a turbine, a compressor, a core shaft, and a core exhaust nozzle, the core exhaust nozzle having a core exhaust nozzle pressure ratio calculated using total pressure at the core nozzle exit; a fan including a plurality of fan blades; and a nacelle surrounding the fan and the engine core and defining a bypass duct, the bypass duct including a bypass exhaust nozzle, the bypass exhaust nozzle having a bypass exhaust nozzle pressure ratio calculated using total pressure at the bypass nozzle exit;
wherein a bypass to core ratio of: bypass ⁢ ⁢ exhaust ⁢ ⁢ nozzle ⁢ ⁢ pressure ⁢ ⁢ ratio core ⁢ ⁢ exhaust ⁢ ⁢ nozzle ⁢ ⁢ pressure ⁢ ⁢ ratio is configured to be in the range from 1.1 to 2.0 under aircraft cruise conditions.