公开(公告)号：US20210164417A1

公开(公告)日：2021-06-03

申请号：US17175092

申请日：2021-02-12

Applicant: ROLLS-ROYCE PLC

Inventor： Richard G STRETTON ,  Michael C WILLMOT

IPC: F02K3/06 ,  F02C3/04 ,  F02C7/04 ,  F02C7/36

Abstract: A gas turbine engine for an aircraft including: engine core including a turbine; and fan including a plurality of fan blades extending radially from a hub, each fan blade having a leading and trailing edge. Turbine includes a lowest pressure turbine stage having a row of rotor blades each extending radially and having a leading and trailing edge. A fan-turbine radius difference is measured as radial distance between: a point on a circle swept by a radially outer tip of the trailing edge of each of the rotor blades of the lowest pressure stage of the turbine; and a point on a circle swept by a radially outer tip of the leading edge of each of fan blades; and a fan speed to fan-turbine radius ratio defined as: the   maximum   take - off   rotational   speed   of   the   fan fan - turbine   radius   difference   ( 120 ) is in a range between 0.8 rpm/mm to 5 rpm/mm.

公开(公告)号：US20210148306A1

公开(公告)日：2021-05-20

申请号：US17146910

申请日：2021-01-12

Applicant: ROLLS-ROYCE plc

Inventor： Richard G. STRETTON ,  Michael C WILLMOT

IPC: F02K3/06 ,  F02C3/04 ,  F02C7/04 ,  F02C7/36

Abstract: A gas turbine engine for an aircraft including: engine core including a turbine; and fan including a plurality of fan blades extending radially from a hub, each fan blade having a leading and trailing edge. Turbine includes a lowest pressure turbine stage having a row of rotor blades each extending radially and having a leading and trailing edge. A fan-turbine radius difference is measured as radial distance between: a point on a circle swept by a radially outer tip of the trailing edge of each of the rotor blades of the lowest pressure stage of the turbine; and a point on a circle swept by a radially outer tip of the leading edge of each of fan blades; and a fan speed to fan-turbine radius ratio defined as: the   maximum   take  -  off   rotational   speed   of   the   fan fan  -  turbine   radius   difference   ( 120 ) is in a range between 0.8 rpm/mm to 5 rpm/mm.

公开(公告)号：US20250154902A1

公开(公告)日：2025-05-15

申请号：US18916983

申请日：2024-10-16

Applicant: ROLLS-ROYCE plc

Inventor： Michael C WILLMOT ,  Richard PEACE

IPC: F02C7/22 ,  F02C7/268 ,  H01M8/04082

Abstract: A propulsion system comprises a propulsive hydrogen-burning gas turbine engine, a first tank storing liquid hydrogen with an ullage and a first fuel line including a fuel pump and a vaporiser, the first fuel line providing gaseous hydrogen to the engine during operation of the system. A second tank storing gaseous hydrogen is coupled by a second fuel line to the first fuel line at a position thereon between the vaporiser and the engine, providing for engine start-up (when the vaporiser is inoperative) and power-boosting during operation of the system. A duct connects gaseous hydrogen within the second tank to the ullage in the first tank in order maintain pressure in the first tank as liquid hydrogen within it is depleted, preventing cavitation of liquid hydrogen within the fuel pump.

公开(公告)号：US20240376847A1

公开(公告)日：2024-11-14

申请号：US18781563

申请日：2024-07-23

Applicant: ROLLS-ROYCE plc

Inventor： Richard G STRETTON ,  Michael C WILLMOT

IPC: F02C7/36 ,  F01D5/14 ,  F02K3/06 ,  F16H1/28

Abstract: A gas turbine engine (10) for an aircraft comprises an engine core (11) comprising a turbine (19), a compressor (14), and a core shaft (26) connecting the turbine to the compressor; a fan (23) located upstream of the engine core (11), the fan comprising a plurality of fan blades (64) extending from a hub (66); and a gearbox (30) that receives an input from the core shaft (26) and outputs drive to the fan (23) so as to drive the fan at a lower rotational speed than the core shaft. The gas turbine engine (10) has an engine length (110) and a gearbox location (112) relative to a forward region of the fan (23), and a gearbox location ratio of: gearbox location/engine length is in a range from 0.19 to 0.45.

公开(公告)号：US20230028367A1

公开(公告)日：2023-01-26

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

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.