Patents by Inventor Nicholas Howarth
Nicholas Howarth 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: 20210172374Abstract: 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.Type: ApplicationFiled: December 7, 2020Publication date: June 10, 2021Inventors: Nicholas HOWARTH, Amarveer S. MANN
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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
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Patent number: 10590856Abstract: 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.Type: GrantFiled: July 13, 2015Date of Patent: March 17, 2020Assignee: ROLLS-ROYCE PLCInventor: Nicholas Howarth
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Patent number: 10550700Abstract: 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: June 11, 2019Date of Patent: February 4, 2020Assignee: ROLLS-ROYCE plcInventors: Nicholas Howarth, Gareth M Armstrong
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Patent number: 10006374Abstract: 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.Type: GrantFiled: April 16, 2015Date of Patent: June 26, 2018Assignee: ROLLS-ROYCE plcInventors: Alasdair Gardner, Arthur Laurence Rowe, Mark David Taylor, Nicholas Howarth
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Patent number: 9631555Abstract: An air intake guide for a jet propulsion power plant for a supersonic aircraft comprises an intake aperture, an intake center body and an intake adjustment device. The intake aperture has an intake lip, an intake center body is positioned within the aperture, and an intake adjustment device is positioned on a radially inwardly facing surface of the air intake guide downstream of the intake lip. The intake adjustment device comprises a flexible panel and an actuator with the actuator being adapted to deflect the flexible panel in a radially inwardly direction so as to reduce a cross-sectional area of the intake aperture and thereby to position a shock wave at the intake lip.Type: GrantFiled: March 21, 2014Date of Patent: April 25, 2017Assignee: ROLLS-ROYCE plcInventor: Nicholas Howarth
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Patent number: 9387923Abstract: A gas turbine engine (10) having an axial flow direction (X) therethrough in use. The gas turbine engine (10) comprises one or more rotor stages each comprising at least one rotor blade (120) having a root portion (122). The gas turbine engine (10) comprises a shroud (122) located upstream of one or more of the rotor stages relative to the axial flow direction (X). The shroud (122) defines a through passageway (128) extending between an inlet (130) and an outlet (132) which comprises a diffuser region (138). The diffuser region (138) is configured to reduce the axial velocity of air exiting the outlet (132) relative to air entering the diffuser portion (138) in use, wherein the outlet (132) is located such that air exiting the outlet (132) is directed substantially to the root portion (122) only of the rotor blades (120).Type: GrantFiled: June 11, 2013Date of Patent: July 12, 2016Assignee: ROLLS-ROYCE plcInventors: Richard Geoffrey Stretton, Nicholas Howarth
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Publication number: 20160040605Abstract: 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.Type: ApplicationFiled: July 13, 2015Publication date: February 11, 2016Inventor: Nicholas HOWARTH
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Publication number: 20150308353Abstract: 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.Type: ApplicationFiled: April 16, 2015Publication date: October 29, 2015Inventors: Alasdair GARDNER, Arthur Laurence ROWE, Mark David TAYLOR, Nicholas HOWARTH
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Patent number: 8967967Abstract: The present disclosure relates to a propfan engine comprising: one or more rotor stages comprising a plurality of rotors; and an outer wall comprising an outer profile, at least a portion of the outer profile defining a substantially circular cross-section, wherein the diameter of the substantially circular cross-section increases in the direction of flow over the outer wall and downstream of a leading edge of the rotors, and the diameter increases at substantially all points defining the circumference of the substantially circular cross-section.Type: GrantFiled: January 26, 2012Date of Patent: March 3, 2015Assignee: Rolls-Royce PLCInventors: Richard G. Stretton, Nicholas Howarth
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Publication number: 20140311580Abstract: An air intake guide for a jet propulsion power plant for a supersonic aircraft comprises an intake aperture, an intake centre body and an intake adjustment device. The intake aperture has an intake lip, an intake centre body is positioned within the aperture, and an intake adjustment device is positioned on a radially inwardly facing surface of the air intake guide downstream of the intake lip. The intake adjustment device comprises a flexible panel and an actuator with the actuator being adapted to deflect the flexible panel in a radially inwardly direction so as to reduce a cross-sectional area of the intake aperture and thereby to position a shock wave at the intake lip.Type: ApplicationFiled: March 21, 2014Publication date: October 23, 2014Applicant: ROLLS-ROYCE PLCInventor: Nicholas HOWARTH
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Publication number: 20140125121Abstract: A method of operating an electrical power generation system on an aircraft. The method includes assessing a required electrical power of the aircraft, assessing whether a first and a second electrical power source are able to provide the required electrical power in combination, assessing a predetermined condition, determining an operating mode of the first and second electrical power sources to match the predetermined condition, and operating the first and second electrical power sources according to the determined operating mode.Type: ApplicationFiled: September 25, 2013Publication date: May 8, 2014Applicant: ROLLS-ROYCE PLCInventors: Huw Llewelyn EDWARDS, Parag VYAS, Malcolm Laurence HILLEL, Alasdair GARDNER, Sean Patrick ELLIS, Nicholas HOWARTH
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Publication number: 20140017067Abstract: A gas turbine engine (10) having an axial flow direction (X) therethrough in use. The gas turbine engine (10) comprises one or more rotor stages each comprising at least one rotor blade (120) having a root portion (122). The gas turbine engine (10) comprises a shroud (122) located upstream of one or more of the rotor stages relative to the axial flow direction (X). The shroud (122) defines a through passageway (128) extending between an inlet (130) and an outlet (132) which comprises a diffuser region (138). The diffuser region (138) is configured to reduce the axial velocity of air exiting the outlet (132) relative to air entering the diffuser portion (138) in use, wherein the outlet (132) is located such that air exiting the outlet (132) is directed substantially to the root portion (122) only of the rotor blades (120).Type: ApplicationFiled: June 11, 2013Publication date: January 16, 2014Inventors: Richard Geoffrey STRETTON, Nicholas HOWARTH
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Publication number: 20130343892Abstract: The present disclosure relates to a propfan engine comprising: one or more rotor stages comprising a plurality of rotors; and an outer wall comprising an outer profile, at least a portion of the outer profile defining a substantially circular cross-section, wherein the diameter of the substantially circular cross-section increases in the direction of flow over the outer wall and downstream of a leading edge of the rotors, and the diameter increases at substantially all points defining the circumference of the substantially circular cross-section.Type: ApplicationFiled: January 26, 2012Publication date: December 26, 2013Applicant: ROLLS-ROYCE PLCInventors: Richard G. STRETTON, Nicholas HOWARTH
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Patent number: 8453458Abstract: Variation in the available mixing plane areas in an exhaust arrangement of a gas turbine engine enables alteration and configuration for better thermal cycle performance of that engine. A shaped centre fairing is associated with an exit nozzle and a bypass duct such that channels between the fairing, nozzle exit and duct can be adjusted to change the available areas. Such variation is achieved by relative axial displacement, typically of the exit nozzle using an appropriate mechanism.Type: GrantFiled: January 23, 2007Date of Patent: June 4, 2013Assignee: Rolls-Royce PLCInventors: John R Whurr, Richard G Stretton, David M Beaven, Nicholas Howarth