Patents by Inventor Gareth M. Armstrong
Gareth M. Armstrong has filed for patents to protect the following inventions. This listing includes patent applications that are pending as well as patents that have already been granted by the United States Patent and Trademark Office (USPTO).
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Publication number: 20230392554Abstract: 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.Type: ApplicationFiled: August 22, 2023Publication date: December 7, 2023Applicant: ROLLS-ROYCE plcInventors: Nicholas HOWARTH, Gareth M ARMSTRONG
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Patent number: 11781491Abstract: 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.Type: GrantFiled: November 15, 2022Date of Patent: October 10, 2023Assignee: ROLLS-ROYCE plcInventors: Nicholas Howarth, Gareth M Armstrong
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Publication number: 20230242264Abstract: 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.Type: ApplicationFiled: January 18, 2023Publication date: August 3, 2023Applicant: ROLLS-ROYCE PLCInventors: Gareth M. ARMSTRONG, Nicholas HOWARTH
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Patent number: 11698030Abstract: 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.Type: GrantFiled: April 28, 2022Date of Patent: July 11, 2023Assignee: ROLLS-ROYCE plcInventors: Nicholas Howarth, Gareth M Armstrong
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Publication number: 20230079630Abstract: 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.Type: ApplicationFiled: November 15, 2022Publication date: March 16, 2023Applicant: ROLLS-ROYCE plcInventors: Nicholas HOWARTH, Gareth M. ARMSTRONG
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Patent number: 11584532Abstract: 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.Type: GrantFiled: December 27, 2021Date of Patent: February 21, 2023Assignee: ROLLS-ROYCE plcInventors: Gareth M Armstrong, Nicholas Howarth
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Publication number: 20220412269Abstract: 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.Type: ApplicationFiled: April 28, 2022Publication date: December 29, 2022Applicant: ROLLS-ROYCE PLCInventors: Nicholas HOWARTH, Gareth M. ARMSTRONG
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Publication number: 20220403743Abstract: 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.Type: ApplicationFiled: August 22, 2022Publication date: December 22, 2022Applicant: ROLLS-ROYCE plcInventors: Nicholas HOWARTH, Gareth M ARMSTRONG
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Patent number: 11459893Abstract: 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.Type: GrantFiled: March 25, 2021Date of Patent: October 4, 2022Assignee: ROLLS-ROYCE PLCInventors: Nicholas Howarth, Gareth M Armstrong
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Patent number: 11346287Abstract: 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.Type: GrantFiled: February 9, 2021Date of Patent: May 31, 2022Assignee: ROLLS-ROYCE PLCInventors: Nicholas Howarth, Gareth M Armstrong
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Publication number: 20220119120Abstract: 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.Type: ApplicationFiled: December 27, 2021Publication date: April 21, 2022Applicant: ROLLS-ROYCE PLCInventors: Gareth M ARMSTRONG, Nicholas HOWARTH
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Patent number: 11242155Abstract: 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.Type: GrantFiled: June 11, 2019Date of Patent: February 8, 2022Assignee: ROLLS-ROYCE plcInventors: Gareth M Armstrong, Nicholas Howarth
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Patent number: 11073019Abstract: 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.Type: GrantFiled: January 15, 2020Date of Patent: July 27, 2021Assignee: ROLLS-ROYCE plcInventors: Gareth M Armstrong, Michael P Keenan
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Publication number: 20210207483Abstract: 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.Type: ApplicationFiled: March 25, 2021Publication date: July 8, 2021Applicant: ROLLS-ROYCE PLCInventors: Nicholas HOWARTH, Gareth M ARMSTRONG
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Publication number: 20210164401Abstract: 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.Type: ApplicationFiled: February 9, 2021Publication date: June 3, 2021Applicant: ROLLS-ROYCE PLCInventors: Nicholas HOWARTH, Gareth M ARMSTRONG
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Patent number: 10982550Abstract: 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.Type: GrantFiled: December 13, 2019Date of Patent: April 20, 2021Assignee: ROLLS-ROYCE plcInventors: Nicholas Howarth, Gareth M Armstrong
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Patent number: 10961916Abstract: 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.Type: GrantFiled: June 18, 2019Date of Patent: March 30, 2021Assignee: ROLLS-ROYCE plcInventors: Nicholas Howarth, Gareth M Armstrong
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Publication number: 20200291865Abstract: 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.Type: ApplicationFiled: June 18, 2019Publication date: September 17, 2020Applicant: ROLLS-ROYCE plcInventors: Nicholas HOWARTH, Gareth M ARMSTRONG
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Publication number: 20200290743Abstract: 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.Type: ApplicationFiled: June 11, 2019Publication date: September 17, 2020Applicant: ROLLS-ROYCE plcInventors: Gareth M ARMSTRONG, Nicholas HOWARTH
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Publication number: 20200291785Abstract: 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.Type: ApplicationFiled: December 13, 2019Publication date: September 17, 2020Applicant: ROLLS-ROYCE plcInventors: Nicholas HOWARTH, Gareth M ARMSTRONG