-
公开(公告)号：US11732603B2
公开(公告)日：2023-08-22
申请号：US17969826
申请日：2022-10-20
Applicant: ROLLS-ROYCE PLC
Inventor： Ian J Bousfield , Duncan A MacDougall
CPC classification number: F01D25/02 , F01D5/141 , F01D5/147 , F04D29/324 , F05D2220/32 , F05D2240/303 , F05D2240/307
Abstract: A gas turbine engine includes a fan mounted to rotate about a main longitudinal axis; an engine core, including a compressor, a combustor, and turbine coupled to the compressor through a shaft, and reduction gearbox; wherein the compressor includes a plurality of stages, each stage including a respective rotor and stator, a first stage of the plurality of stages being arranged at an inlet and including a first rotor with a plurality of blades; each blade extending chordwise from a leading edge to a trailing edge, and from root to tip for a span height, wherein 0% of the span height corresponds to the root and 100% of span height corresponds to tip; wherein a ratio of a leading edge radius of each of the plurality of first rotor blades at 0% span height to a minimum leading edge radius is comprised between 1 and 1.50.
-
公开(公告)号：US11512607B2
公开(公告)日：2022-11-29
申请号：US17235234
申请日：2021-04-20
Applicant: ROLLS-ROYCE PLC
Inventor： Ian J Bousfield , Duncan A MacDougall
Abstract: A gas turbine engine includes a fan mounted to rotate about a main longitudinal axis; an engine core, including a compressor, a combustor, and turbine coupled to the compressor through a shaft; and reduction gearbox; wherein the compressor includes a plurality of stages, each stage including a respective rotor and stator, a first stage of the plurality of stages being arranged at an inlet and including a first rotor with a plurality of blades; each blade extending chordwise from a leading edge to a trailing edge, and from root to tip for a span height, wherein 0% of the span height corresponds to the root and 100% of span height corresponds to tip; wherein a ratio of a leading edge radius of each of the plurality of first rotor blades at 0% span height to a minimum leading edge radius is comprised between 1 and 1.50.
-
公开(公告)号：US10995677B2
公开(公告)日：2021-05-04
申请号：US16813830
申请日：2020-03-10
Applicant: ROLLS-ROYCE PLC
Inventor： Ian J Bousfield , Duncan A Macdougall
Abstract: A gas turbine engine comprises a fan mounted to rotate about a main longitudinal axis; an engine core, comprising in axial flow series a compressor, a combustor, and a turbine coupled to the compressor through a shaft; a reduction gearbox that receives an input from the shaft and outputs drive to the fan so as to drive the fan at a lower rotational speed than the shaft; wherein the compressor comprises a first stage at an inlet and a second stage, downstream of the first stage, comprising respectively a first rotor with a row of first blades and a second rotor with a row of second blades, the first and second blades comprising respective leading edges, trailing edges and tips, and wherein the ratio of a maximum leading edge radius of the first blades to a maximum leading edge radius of the second blades is greater than 2.8.
-
公开(公告)号：US11686248B2
公开(公告)日：2023-06-27
申请号：US17333631
申请日：2021-05-28
Applicant: ROLLS-ROYCE PLC
Inventor： Ian J Bousfield , Duncan A MacDougall
CPC classification number: F02C7/05 , F01D5/143 , F01D21/10 , F01D25/02 , F02C7/04 , F02C7/30 , F02K1/82
Abstract: A core duct assembly for a gas turbine engine includes a core duct including an outer and an inner wall, the outer wall having an interior surface; a gas flow path member extending across the gas flow path at least partly between the inner and outer walls, the rotor blade having a radial span extending from a blade platform to a blade tip, wherein an upstream wall axis is defined as an axis tangential to a point on a first portion of the interior surface of the outer wall of the core duct extending downstream from the gas flow path member, the upstream wall axis lying in a longitudinal plane of the gas turbine engine containing the rotational axis of the engine, and wherein the upstream wall axis intersects the rotor blade at a point spaced radially inward from the blade tip of the rotor blade.
-
公开(公告)号：US11519363B2
公开(公告)日：2022-12-06
申请号：US17196382
申请日：2021-03-09
Applicant: ROLLS-ROYCE plc
Inventor： Michael O Hales , Craig W Bemment , Benjamin J Sellers , Ian J Bousfield , Amarveer S Mann
Abstract: A gas turbine engine (10) comprising: a high pressure turbine (17); a low pressure turbine (19); a high pressure compressor (15) coupled to the high pressure turbine (17) by a high pressure shaft (27); a propulsor (23) and a low pressure compressor (14) coupled to the low pressure turbine (19) via a low pressure shaft (26) and a reduction gearbox (30); wherein the low pressure compressor (14) consists of four compressor stages (14) and defines a cruise pressure ratio of between 2.4:1 and 3.3:1; the high pressure compressor (15) defines a cruise pressure ratio of less than 17:1; and the high pressure compressor (15) and low pressure compressor (14) together define a cruise core overall pressure ratio of greater than 36:1.
-
Patent Agency Ranking
公开(公告)号：US11480134B2
公开(公告)日：2022-10-25
申请号：US17196546
申请日：2021-03-09
Applicant: ROLLS-ROYCE PLC
Inventor： Ian J Bousfield , Michael O Hales , Rory D Stieger
Abstract: A gas turbine engine including: a high pressure turbine, a low pressure turbine, a high pressure compressor coupled to the high pressure turbine by a high pressure shaft, a propulsor and a low pressure compressor coupled to the low pressure turbine via a low pressure shaft and a reduction gearbox; wherein the high pressure compressor defines an average stage pressure ratio at cruise conditions of between 1.25 and 1.35 and consists of 10 or 11 stages; and the high pressure compressor and low pressure compressor together define a core overall pressure ratio at cruise conditions of between 40:1 and 60:1.
-
公开(公告)号：US12286895B2
公开(公告)日：2025-04-29
申请号：US18825241
申请日：2024-09-05
Applicant: ROLLS-ROYCE plc
Inventor： Ian J Bousfield , Duncan A Macdougall
Abstract: A gas turbine engine comprises a fan mounted to rotate about a main longitudinal axis; an engine core, comprising in axial flow series a compressor, a combustor, and a turbine coupled to the compressor through a shaft; a reduction gearbox that receives an input from the shaft and outputs drive to the fan so as to drive the fan at a lower rotational speed than the shaft; wherein the compressor comprises a first stage at an inlet and a second stage, downstream of the first stage, comprising respectively a first rotor with a row of first blades and a second rotor with a row of second blades, the first and second blades comprising respective leading edges, trailing edges and tips, and wherein the ratio of a maximum leading edge radius of the first blades to a maximum leading edge radius of the second blades is greater than 2.8.
-
公开(公告)号：US11629668B2
公开(公告)日：2023-04-18
申请号：US17196546
申请日：2021-03-09
Applicant: ROLLS-ROYCE PLC
Inventor： Ian J Bousfield , Michael O Hales , Rory D Stieger
Abstract: A gas turbine engine including: a high pressure turbine, a low pressure turbine, a high pressure compressor coupled to the high pressure turbine by a high pressure shaft, a propulsor and a low pressure compressor coupled to the low pressure turbine via a low pressure shaft and a reduction gearbox; wherein the high pressure compressor defines an average stage pressure ratio at cruise conditions of between 1.25 and 1.35 and consists of 10 or 11 stages; and the high pressure compressor and low pressure compressor together define a core overall pressure ratio at cruise conditions of between 40:1 and 60:1.
-
公开(公告)号：US11619135B2
公开(公告)日：2023-04-04
申请号：US17218617
申请日：2021-03-31
Applicant: ROLLS-ROYCE PLC
Inventor： Ian J Bousfield , Duncan A MacDougall
Abstract: A gas turbine engine comprises a fan mounted to rotate about a main longitudinal axis; an engine core, comprising in axial flow series a compressor, a combustor, and a turbine coupled to the compressor through a shaft; a reduction gearbox that receives an input from the shaft and outputs drive to the fan so as to drive the fan at a lower rotational speed than the shaft; wherein the compressor comprises a first stage at an inlet and a second stage, downstream of the first stage, comprising respectively a first rotor with a row of first blades and a second rotor with a row of second blades, the first and second blades comprising respective leading edges, trailing edges and tips, and wherein the ratio of a maximum leading edge radius of the first blades to a maximum leading edge radius of the second blades is greater than 2.8.
-
公开(公告)号：US11047311B2
公开(公告)日：2021-06-29
申请号：US16437283
申请日：2019-06-11
Applicant: ROLLS-ROYCE plc
Inventor： Ian J Bousfield , Duncan A MacDougall
Abstract: A core duct assembly for a gas turbine engine, the core duct assembly including: a core duct including an outer and an inner wall, the outer wall having an interior surface; a gas flow path member extending across the gas flow path at least partly between the inner and outer walls, the rotor blade having a radial span extending from a blade platform to a blade tip, wherein an upstream wall axis is defined as an axis tangential to a point on a first portion of the interior surface of the outer wall of the core duct extending downstream from the gas flow path member, the upstream wall axis lying in a longitudinal plane of the gas turbine engine containing the rotational axis of the engine, and wherein the upstream wall axis intersects the rotor blade at a point spaced radially inward from the blade tip of the rotor blade.
-
-
-
-
-
-
-
-
-
Search Results
13
records
(took 0.015 seconds)
