公开(公告)号：US10982550B2

公开(公告)日：2021-04-20

申请号：US16713666

申请日：2019-12-13

Applicant: ROLLS-ROYCE plc

Inventor： Nicholas Howarth ,  Gareth M Armstrong

IPC: F01D5/14 ,  F02K3/06 ,  F02C3/02 ,  F02C7/36 ,  F04D29/68 ,  F04D19/02 ,  F01D5/28 ,  F02C3/04 ,  F02C3/107 ,  F02C7/04 ,  F02K3/068

Abstract: A gas turbine engine that has improved fuel burn provides operability and/or maintenance requirements when installed on an aircraft. The gas turbine engine is provided with a core compressor that includes twelve, thirteen or fourteen rotor stages. The gas turbine engine has a ratio of a core compressor aspect ratio divided by a core compressor pressure ratio is in the range of from 0.03 to 0.09. This results in an optimum balance between installation benefits, operability, maintenance requirements and engine efficiency when the gas turbine engine is installed on an aircraft.

公开(公告)号：US10590856B2

公开(公告)日：2020-03-17

申请号：US14797871

申请日：2015-07-13

Applicant: ROLLS-ROYCE PLC

Inventor： Nicholas Howarth

IPC: F02C9/18 ,  F01D17/10 ,  F02C3/13 ,  F01D25/26 ,  F02C3/04

Abstract: A gas turbine engine including a compressor, a turbine having one or more stages and a combustor, the combustor being located between the compressor and turbine. The gas turbine engine further includes a bleed from a core defined by a core duct, the core duct surrounding and extending between the turbine and combustor at least. The bleed includes at least one inlet located downstream of the combustor and upstream of at least one of the turbine stages. The turbine is arranged in use to drive the compressor. The bleed is arranged to be controllable in use to selectively bleed air from the core through the inlet and to thereby control the power delivered by the turbine to the compressor.

公开(公告)号：US10550700B1

公开(公告)日：2020-02-04

申请号：US16437356

申请日：2019-06-11

Applicant: ROLLS-ROYCE plc

Inventor： Nicholas Howarth ,  Gareth M Armstrong

IPC: F02C7/36 ,  F02C3/107 ,  F02K3/06 ,  F01D5/14 ,  F02C3/02 ,  F04D29/68 ,  F04D19/02 ,  F01D5/28 ,  F02C3/04 ,  F02C7/04 ,  F02K3/068

Abstract: A gas turbine engine that has improved fuel burn provides operability and/or maintenance requirements when installed on an aircraft. The gas turbine engine is provided with a core compressor that includes twelve, thirteen or fourteen rotor stages. The gas turbine engine has a ratio of a core compressor aspect ratio divided by a core compressor pressure ratio is in the range of from 0.03 to 0.09. This results in an optimum balance between installation benefits, operability, maintenance requirements and engine efficiency when the gas turbine engine is installed on an aircraft.

公开(公告)号：US10006374B2

公开(公告)日：2018-06-26

申请号：US14688423

申请日：2015-04-16

Applicant: ROLLS-ROYCE PLC

Inventor： Alasdair Gardner ,  Arthur Laurence Rowe ,  Mark David Taylor ,  Nicholas Howarth

IPC: F02C9/22 ,  F01D17/16 ,  F02C9/20 ,  F02C9/54 ,  F02C6/00 ,  F02C7/36

CPC classification number: F02C9/22 ,  F01D17/162 ,  F02C6/00 ,  F02C7/36 ,  F02C9/20 ,  F02C9/54 ,  F05D2220/32 ,  F05D2220/76 ,  F05D2270/05 ,  F05D2270/06 ,  F05D2270/303 ,  F05D2270/304 ,  Y02T50/671

Abstract: An engine that has, in axial flow series, booster compressor, core compressor, combustion equipment, core turbine, and low-pressure turbine. Core turbine drives core compressor via an interconnecting high-pressure shaft. Low-pressure turbine drives booster compressor via an interconnecting low-pressure shaft. Low-pressure turbine also drives external load having a defined speed characteristic that dictates speed of the low-pressure turbine and booster compressor. Booster compressor has one or more rows of variable stator vanes. The method includes: scheduling variation in the angle of variable stator vanes as a function of speed of the booster compressor wherein the vanes open as booster compressor speed increases; measuring or setting one or more operational parameters which are determinative of temperature at entry to core turbine; and biasing scheduling of angle variation of variable stator vanes as a function of operational parameter(s) to reduce variation in temperature at entry to core turbine.

公开(公告)号：US12071901B2

公开(公告)日：2024-08-27

申请号：US18236666

申请日：2023-08-22

Applicant: ROLLS-ROYCE plc

Inventor： Nicholas Howarth ,  Gareth M Armstrong

IPC: F02C7/36 ,  F01D15/12 ,  F02C3/113

CPC classification number: F02C7/36 ,  F01D15/12 ,  F02C3/113

Abstract: A gas turbine engine has a compression system blade ratio defined as the ratio of the height of a fan blade to the height of the most downstream compressor blade in the range of from 45 to 95. This results in an optimum balance between installation benefits, operability, maintenance requirements and engine efficiency when the gas turbine engine is installed on an aircraft.

公开(公告)号：US12024301B2

公开(公告)日：2024-07-02

申请号：US18098463

申请日：2023-01-18

Applicant: ROLLS-ROYCE PLC

Inventor： Gareth M Armstrong ,  Nicholas Howarth

IPC: B64D27/10 ,  B64D33/02 ,  F01D5/28 ,  F02C3/04 ,  F02C3/06 ,  F02C3/107 ,  F02C7/04 ,  F02K3/06 ,  F02K3/068

CPC classification number: B64D27/10 ,  F01D5/28 ,  F02C3/04 ,  F02C3/06 ,  F02C3/107 ,  F02C7/04 ,  F02K3/06 ,  F02K3/068 ,  B64D2033/0286 ,  F05D2220/3215 ,  F05D2220/3219 ,  F05D2220/36 ,  F05D2260/40311

Abstract: A gas turbine engine has a compression system radius ratio defined as the ratio of the radius of the tip of a fan blade to the radius of the tip of the most downstream compressor blade in the range of from 5 to 9. This results in an optimum balance between installation benefits, operability, maintenance requirements and engine efficiency when the gas turbine engine is installed on an aircraft.

公开(公告)号：US12006835B2

公开(公告)日：2024-06-11

申请号：US17892696

申请日：2022-08-22

Applicant: ROLLS-ROYCE plc

Inventor： Nicholas Howarth ,  Gareth M Armstrong

IPC: F01D5/14 ,  F01D5/28 ,  F02C3/02 ,  F02C3/04 ,  F02C3/107 ,  F02C7/04 ,  F02C7/36 ,  F02K3/06 ,  F02K3/068 ,  F04D19/02 ,  F04D29/68

CPC classification number: F01D5/145 ,  F01D5/141 ,  F01D5/28 ,  F02C3/02 ,  F02C3/04 ,  F02C3/107 ,  F02C7/04 ,  F02C7/36 ,  F02K3/06 ,  F02K3/068 ,  F04D19/024 ,  F04D29/681

Abstract: A gas turbine engine that has improved fuel burn provides operability and/or maintenance requirements when installed on an aircraft. The gas turbine engine is provided with a core compressor that includes twelve, thirteen or fourteen rotor stages. The gas turbine engine has a ratio of a core compressor aspect ratio divided by a core compressor pressure ratio is in the range of from 0.03 to 0.09. This results in an optimum balance between installation benefits, operability, maintenance requirements and engine efficiency when the gas turbine engine is installed on an aircraft.

公开(公告)号：US11698030B2

公开(公告)日：2023-07-11

申请号：US17731877

申请日：2022-04-28

Applicant: ROLLS-ROYCE PLC

Inventor： Nicholas Howarth ,  Gareth M Armstrong

IPC: F01D15/12 ,  F02C7/36 ,  F02C3/113

CPC classification number: F02C7/36 ,  F01D15/12 ,  F02C3/113

Abstract: A gas turbine engine has a compression system blade ratio defined as the ratio of the height of a fan blade to the height of the most downstream compressor blade in the range of from 45 to 95. This results in an optimum balance between installation benefits, operability, maintenance requirements and engine efficiency when the gas turbine engine is installed on an aircraft.

公开(公告)号：US11339717B2

公开(公告)日：2022-05-24

申请号：US17113910

申请日：2020-12-07

Applicant: ROLLS-ROYCE plc

Inventor： Nicholas Howarth ,  Amarveer S Mann

IPC: B64D13/08 ,  F02C6/08 ,  F02C7/18 ,  B64D13/06 ,  F02C7/057 ,  F02C7/04 ,  F02C7/36

Abstract: The present disclosure relates to a gas turbine engine for an aircraft comprising: an engine core comprising a turbine, a compressor, and a core shaft; and an environmental control system mounted on the engine core comprising a first air passage arranged to deliver air from outside the engine core to an aircraft cabin and/or for wing anti icing, a subsidiary compressor located in the first air passage and arranged to compress air in the first air passage, the subsidiary compressor being powered by the core shaft, and a second air passage arranged to inject air from the compressor into the first air passage.

公开(公告)号：US10961916B2

公开(公告)日：2021-03-30

申请号：US16443938

申请日：2019-06-18

Applicant: ROLLS-ROYCE plc

Inventor： Nicholas Howarth ,  Gareth M Armstrong

IPC: F02C7/36 ,  F02C3/113 ,  F01D15/12

Abstract: A gas turbine engine has a compression system blade ratio defined as the ratio of the height of a fan blade to the height of the most downstream compressor blade in the range of from 45 to 95. This results in an optimum balance between installation benefits, operability, maintenance requirements and engine efficiency when the gas turbine engine is installed on an aircraft.