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公开(公告)号：US11346287B2
公开(公告)日：2022-05-31
申请号：US17171439
申请日：2021-02-09
Applicant: ROLLS-ROYCE PLC
Inventor： Nicholas Howarth , Gareth M Armstrong
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.
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公开(公告)号：US11781491B2
公开(公告)日：2023-10-10
申请号：US17987516
申请日：2022-11-15
Applicant: ROLLS-ROYCE plc
Inventor： Nicholas Howarth , Gareth M Armstrong
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.
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公开(公告)号：US11584532B2
公开(公告)日：2023-02-21
申请号：US17562609
申请日：2021-12-27
Applicant: ROLLS-ROYCE PLC
Inventor： Gareth M Armstrong , Nicholas Howarth
IPC: B64D27/10 , F02C3/06 , F02K3/06 , F02K3/068 , F02C3/04 , F02C7/04 , B64D33/02 , F01D5/28 , F02C3/107
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.
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公开(公告)号：US11073019B2
公开(公告)日：2021-07-27
申请号：US16743311
申请日：2020-01-15
Applicant: ROLLS-ROYCE plc
Inventor： Gareth M Armstrong , Michael P Keenan
IPC: F01D5/02
Abstract: The present disclosure relates to a metallic shaft for connecting components of a gas turbine engine. Example embodiments include a metallic shaft (400) for connecting components of a gas turbine engine, the shaft (400) having a longitudinal axis (410) and comprising: a first section (401) extending from a first end (403) of the shaft (400) to a joint (405), the first section (401) composed of a material having a first thermal expansion coefficient along the longitudinal axis (410); a second section (402) extending from a second opposing end (404) of the shaft to the joint (405), the second section (402) composed of a material having a second thermal expansion coefficient along the longitudinal axis (410) that is different to the first thermal expansion coefficient.
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公开(公告)号：US11459893B2
公开(公告)日：2022-10-04
申请号：US17212554
申请日：2021-03-25
Applicant: ROLLS-ROYCE PLC
Inventor： Nicholas Howarth , Gareth M Armstrong
IPC: F01D5/14 , F02C7/04 , F02C3/02 , F02C7/36 , F04D29/68 , F04D19/02 , F01D5/28 , F02C3/04 , F02C3/107 , F02K3/06 , 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.
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Patent Agency Ranking
公开(公告)号：US11242155B2
公开(公告)日：2022-02-08
申请号：US16437284
申请日：2019-06-11
Applicant: ROLLS-ROYCE plc
Inventor： Gareth M Armstrong , Nicholas Howarth
IPC: B64D27/10 , F02C3/06 , F02K3/06 , F02K3/068 , F02C3/04 , F01D5/28 , F02C7/04 , F02C3/107 , B64D33/02
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.
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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.
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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.
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公开(公告)号：US20250012221A1
公开(公告)日：2025-01-09
申请号：US18785906
申请日：2024-07-26
Applicant: ROLLS-ROYCE plc
Inventor： Nicholas HOWARTH , Gareth M Armstrong
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.
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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
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.
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