公开(公告)号：US20250154894A1

公开(公告)日：2025-05-15

申请号：US18916958

申请日：2024-10-16

Applicant: Rolls-Royce plc

Inventor： Michael C. WILLMOT ,  Gergana Yanakieva DIMITROVA

IPC: F02C3/22 ,  F02C7/08

Abstract: A propulsion system comprises a propulsive hydrogen-burning gas turbine engine, a fuel cell stack auxiliary power unit (APU) and a first tank arranged to store liquid hydrogen with an ullage. A first fuel line includes a first pump and a first vaporiser and transports hydrogen from the first tank to combustion apparatus of the engine during operation of the propulsion system. A second fuel line includes a second fuel pump and a second vaporiser and transports hydrogen from the first tank to the fuel cell stack APU. A duct connects the second fuel line at a position thereon between the second vaporiser and the fuel cell stack to the ullage of the first tank, providing for pressure in the first tank to be maintained therein as liquid hydrogen within the first tank is depleted, thus avoiding cavitation of liquid hydrogen within the first fuel pump.

公开(公告)号：US20220056916A1

公开(公告)日：2022-02-24

申请号：US17520126

申请日：2021-11-05

Applicant: ROLLS-ROYCE plc

Inventor： Richard G. STRETTON ,  Michael C. WILLMOT

IPC: F04D19/00 ,  F01D5/12

Abstract: A gas turbine engine for an aircraft includes an engine core having a core length and comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor, the turbine comprising a lowest pressure rotor stage, the turbine having a turbine diameter at the lowest pressure rotor stage; and a fan located upstream of the engine core, the fan comprising a plurality of fan blades extending from a hub, the hub and fan blades together defining a fan face having a fan face area and a fan tip radius, wherein a ratio of the fan tip radius to the turbine diameter at the lowest pressure rotor stage is in a range from 1.2 to 2.0; and wherein the engine core length is in a range from 150 cm to 320 cm.

公开(公告)号：US20200347848A1

公开(公告)日：2020-11-05

申请号：US16825504

申请日：2020-03-20

Applicant: ROLLS-ROYCE plc

Inventor： Richard G. STRETTON ,  Michael C. WILLMOT

IPC: F04D19/00 ,  F01D5/12

Abstract: A gas turbine engine for an aircraft includes an engine core including a turbine, compressor, and core shaft connecting turbine to compressor; a fan located upstream of the engine core and including a plurality of fan blades each having a leading and trailing edge. The turbine includes a lowest pressure turbine stage having a row of rotor blades, each rotor blades extending radially and having a leading and trailing edge. The engine has a fan tip axis that joins a radially outer tip of the leading edge of a fan blade and the radially outer tip of the trailing edge of a rotor blade of the lowest pressure stage. The fan tip axis lies in a longitudinal plane which contains a centreline of engine. A fan axis angle is defined as the angle between fan tip axis and centreline, and is in a range between 10 and 20 degrees.

公开(公告)号：US20200248699A1

公开(公告)日：2020-08-06

申请号：US16825361

申请日：2020-03-20

Applicant: ROLLS-ROYCE plc

Inventor： Richard G. STRETTON ,  Michael C. WILLMOT

IPC: F04D19/00 ,  F01D5/12

Abstract: A gas turbine engine for an aircraft includes an engine core including a turbine, compressor, and core shaft connecting turbine to compressor; a fan located upstream of the engine core and including a plurality of fan blades each having a leading and trailing edge. The turbine includes a lowest pressure turbine stage having a row of rotor blades, each rotor blades extending radially and having a leading and trailing edge. The engine has a fan tip axis that joins a radially outer tip of the leading edge of a fan blade and the radially outer tip of the trailing edge of a rotor blade of the lowest pressure stage. The fan tip axis lies in a longitudinal plane which contains a centreline of engine. A fan axis angle is defined as the angle between fan tip axis and centreline, and is in a range between 10 and 20 degrees.

公开(公告)号：US20210301827A1

公开(公告)日：2021-09-30

申请号：US17338159

申请日：2021-06-03

Applicant: ROLLS-ROYCE plc

Inventor： Richard G. STRETTON ,  Michael C. WILLMOT

IPC: F04D19/00 ,  F01D5/12

Abstract: A gas turbine engine for an aircraft includes an engine core including an engine core, a turbine, a compressor, and a core shaft connecting the turbine to the compressor, the turbine comprising a lowest pressure rotor stage, the turbine having a turbine diameter at the lowest pressure rotor stage. A fan located upstream of the engine core, the fan comprising a plurality of fan blades extending from a hub, the hub and fan blades together defining a fan face having a fan face area and a fan tip radius. A ratio of the fan tip radius to the turbine diameter at the lowest pressure rotor stage is in the range from 1.2 to 2.0.

公开(公告)号：US20210164478A1

公开(公告)日：2021-06-03

申请号：US17174967

申请日：2021-02-12

Applicant: ROLLS-ROYCE plc

Inventor： Richard G. STRETTON ,  Michael C. WILLMOT

IPC: F04D19/00 ,  F01D5/12

Abstract: A gas turbine engine for an aircraft includes an engine core including a turbine, compressor, and core shaft connecting turbine to compressor; a fan located upstream of the engine core and including a plurality of fan blades each having a leading and trailing edge. The turbine includes a lowest pressure turbine stage having a row of rotor blades, each rotor blades extending radially and having a leading and trailing edge. The engine has a fan tip axis that joins a radially outer tip of the leading edge of a fan blade and the radially outer tip of the trailing edge of a rotor blade of the lowest pressure stage. The fan tip axis lies in a longitudinal plane which contains a centreline of engine. A fan axis angle is defined as the angle between fan tip axis and centreline, and is in a range between 10 and 20 degrees.

公开(公告)号：US20200347803A1

公开(公告)日：2020-11-05

申请号：US16929806

申请日：2020-07-15

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.