Patents by Inventor Matthew R. Feulner
Matthew R. Feulner 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: 20230323818Abstract: A fan section includes a fan with fan blades. The fan section drives air along a bypass flow path in a bypass duct. A gear reduction is in driving engagement with the fan and has a gear reduction ratio of greater than 3.0 and less than 4.0. A low spool includes a low pressure turbine that drives a low pressure compressor and drives the gear reduction to drive the fan at a speed slower than the low pressure turbine. The low pressure compressor is a four-stage low pressure compressor. The low pressure turbine is a three-stage low pressure turbine. A high spool including a high pressure turbine that drives a high pressure compressor. The high pressure compressor is a nine-stage high pressure compressor. The high pressure turbine is a two-stage high pressure turbine. An exhaust gas exit temperature of greater than 900 degrees Fahrenheit and less than 1000 degrees Fahrenheit at maximum take-off.Type: ApplicationFiled: May 26, 2023Publication date: October 12, 2023Inventors: Stephen G. Pixton, Matthew R. Feulner, Marc J. Muldoon, Xinwen Xiao
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Patent number: 11754000Abstract: A fan section includes a fan with fan blades. The fan section drives air along a bypass flow path in a bypass duct. A gear reduction is in driving engagement with the fan and has a gear reduction ratio of greater than 3.0 and less than 4.0. A low spool includes a low pressure turbine that drives a low pressure compressor and drives the gear reduction to drive the fan at a speed slower than the low pressure turbine. The low pressure compressor is a four-stage low pressure compressor. The low pressure turbine is a three-stage low pressure turbine. A high spool including a high pressure turbine that drives a high pressure compressor. The high pressure compressor is a nine-stage high pressure compressor. The high pressure turbine is a two-stage high pressure turbine. An exhaust gas exit temperature of greater than 900 degrees Fahrenheit and less than 1000 degrees Fahrenheit at maximum take-off.Type: GrantFiled: July 19, 2021Date of Patent: September 12, 2023Assignee: RTX CorporationInventors: Stephen G. Pixton, Matthew R. Feulner, Marc J. Muldoon, Xinwen Xiao
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Publication number: 20230258192Abstract: A gas turbine engine includes a fan section that includes a fan with fan blades. The fan section drives air along a bypass flow path in a bypass duct. A gear reduction is in driving engagement with the fan and has a gear reduction ratio of greater than 3.0 and less than 4.0. A low spool includes a low pressure turbine that drives a low pressure compressor and drives the gear reduction to drive the fan at a speed slower than the low pressure turbine. The low pressure compressor includes a greater number of stages than the low pressure turbine. A high spool includes a high pressure turbine that drives a high pressure compressor. The high pressure compressor is a nine stage high pressure compressor and includes a pressure ratio per stage of greater than or equal to 1.20 and less than or equal to 1.33.Type: ApplicationFiled: April 28, 2023Publication date: August 17, 2023Inventors: Stephen G. Pixton, Matthew R. Feulner, Marc J. Muldoon, Xinwen Xiao
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Patent number: 11719245Abstract: A gas turbine engine includes a fan section that includes a fan with fan blades. The fan section drives air along a bypass flow path in a bypass duct. A gear reduction is in driving engagement with the fan and has a gear reduction ratio of greater than 3.0 and less than 4.0. A low spool includes a low pressure turbine that drives a low pressure compressor and drives the gear reduction to drive the fan at a speed slower than the low pressure turbine. The low pressure compressor includes a greater number of stages than the low pressure turbine. A high spool includes a high pressure turbine that drives a high pressure compressor. The high pressure compressor is a nine stage high pressure compressor and includes a pressure ratio per stage of greater than or equal to 1.20 and less than or equal to 1.33.Type: GrantFiled: July 19, 2021Date of Patent: August 8, 2023Assignee: Raytheon Technologies CorporationInventors: Stephen G. Pixton, Matthew R. Feulner, Marc J. Muldoon, Xinwen Xiao
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Patent number: 11674407Abstract: A system for controlling blade tip clearances in a gas turbine engine may comprise an active clearance control system and a controller in operable communication with the active clearance control system. The controller may be configured to identify a cruise condition, reduce a thrust limit of the gas turbine engine to a de-rated maximum climb thrust, determine a first target tip clearance based on the de-rated maximum climb thrust, and send a command signal correlating to the first target tip clearance to the active clearance control system.Type: GrantFiled: March 7, 2022Date of Patent: June 13, 2023Assignee: Raytheon Technologies CorporationInventor: Matthew R. Feulner
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Publication number: 20230026997Abstract: A gas turbine engine includes a fan section that includes a fan with fan blades. The fan section drives air along a bypass flow path in a bypass duct. A gear reduction is in driving engagement with the fan and has a gear reduction ratio of greater than 3.0 and less than 4.0. A low spool includes a low pressure turbine that drives a low pressure compressor and drives the gear reduction to drive the fan at a speed slower than the low pressure turbine. A high spool includes a high pressure turbine that drives a high pressure compressor. The high pressure compressor includes a pressure ratio of greater than 6.5 and less than 11.5. A ratio of a product of a pressure ratio of the fan with a pressure ratio of the low pressure compressor pressure to the pressure ratio of the high pressure compressor is greater than 0.35 and less than 0.90. An exhaust gas exit temperature is greater than 900 degrees Fahrenheit and less than 1000 degrees Fahrenheit at maximum take-off.Type: ApplicationFiled: July 19, 2021Publication date: January 26, 2023Inventors: Stephen G. Pixton, Ronald S. Walther, Matthew R. Feulner, Fuhua Ma, Ozhan Turgut
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Publication number: 20230028763Abstract: A gas turbine engine includes a fan section that includes a fan with fan blades. The fan section drives air along a bypass flow path in a bypass duct. A gear reduction is in driving engagement with the fan and has a gear reduction ratio of greater than 3.0 and less than 4.0. A low spool includes a low pressure turbine that drives a low pressure compressor and drives the gear reduction to drive the fan at a speed slower than the low pressure turbine. The low pressure compressor includes a greater number of stages than the low pressure turbine. A high spool includes a high pressure turbine that drives a high pressure compressor. The high pressure compressor is a nine stage high pressure compressor and includes a pressure ratio per stage of greater than or equal to 1.20 and less than or equal to 1.33.Type: ApplicationFiled: July 19, 2021Publication date: January 26, 2023Inventors: Stephen G. Pixton, Matthew R. Feulner, Marc J. Muldoon, Xinwen Xiao
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Publication number: 20230027726Abstract: A gas turbine engine includes a fan section that includes a fan with fan blades. The fan section drives air along a bypass flow path in a bypass duct. A gear reduction is in driving engagement with the fan and has a gear reduction ratio of greater than 3.0 and less than 4.0. A low spool includes a low pressure turbine that drives a low pressure compressor and drives the gear reduction to drive the fan at a speed slower than the low pressure turbine. The low pressure compressor is a five-stage low pressure compressor. The low pressure turbine is four-stage low pressure turbine. A high spool includes a high pressure turbine that drives a high pressure compressor. The high pressure compressor is a nine-stage high pressure compressor. The high pressure turbine is a two-stage high pressure turbine.Type: ApplicationFiled: July 19, 2021Publication date: January 26, 2023Inventors: Stephen G. Pixton, Matthew R. Feulner, Marc J. Muldoon, Xinwen Xiao
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Publication number: 20230022122Abstract: A fan section includes a fan with fan blades. The fan section drives air along a bypass flow path in a bypass duct. A gear reduction is in driving engagement with the fan and has a gear reduction ratio of greater than 3.0 and less than 4.0. A low spool includes a low pressure turbine that drives a low pressure compressor and drives the gear reduction to drive the fan at a speed slower than the low pressure turbine. The low pressure compressor is a four-stage low pressure compressor. The low pressure turbine is a three-stage low pressure turbine. A high spool including a high pressure turbine that drives a high pressure compressor. The high pressure compressor is a nine-stage high pressure compressor. The high pressure turbine is a two-stage high pressure turbine. An exhaust gas exit temperature of greater than 900 degrees Fahrenheit and less than 1000 degrees Fahrenheit at maximum take-off.Type: ApplicationFiled: July 19, 2021Publication date: January 26, 2023Inventors: Stephen G. Pixton, Matthew R. Feulner, Marc J. Muldoon, Xinwen Xiao
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Publication number: 20220341340Abstract: A system for controlling blade tip clearances in a gas turbine engine may comprise an active clearance control system and a controller in operable communication with the active clearance control system. The controller may be configured to identify a cruise condition, reduce a thrust limit of the gas turbine engine to a de-rated maximum climb thrust, determine a first target tip clearance based on the de-rated maximum climb thrust, and send a command signal correlating to the first target tip clearance to the active clearance control system.Type: ApplicationFiled: March 7, 2022Publication date: October 27, 2022Applicant: RAYTHEON TECHNOLOGIES CORPORATIONInventor: Matthew R. Feulner
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Patent number: 11473510Abstract: A method of operating a gas turbine engine includes commanding an acceleration of the gas turbine engine and moving a variable pitch high pressure compressor vane toward an open position thereby reducing an acceleration rate of a high pressure turbine rotor thereby reducing a change in a clearance gap between the high pressure turbine rotor and a blade outer airseal. An active clearance control system of a gas turbine engine includes an engine control system configured to command an acceleration of the gas turbine engine and move a variable pitch high pressure compressor vane toward an open position thereby slowing an acceleration rate of a high pressure turbine rotor thereby reducing a change in a clearance gap between the high pressure turbine rotor and a blade outer airseal located radially outboard of the high pressure turbine rotor.Type: GrantFiled: April 18, 2019Date of Patent: October 18, 2022Assignee: RAYTHEON TECHNOLOGIES CORPORATIONInventors: Joseph Kehoe, Richard P. Meisner, Manuj Dhingra, Patrick D. Couture, Matthew R. Feulner, Brenda J. Lisitano, Christopher L. Ho
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Patent number: 11459946Abstract: A system includes a gas turbine engine having a low speed spool and a high speed spool. The system also includes a spool coupling system configured to mechanically link the low speed spool and the high speed spool. A controller is operable to determine a mode of operation of the gas turbine engine, monitor for a spool coupling activation condition associated with the mode of operation, and activate the spool coupling system based on the controller detecting the spool coupling activation condition. Engagement and power transfer between the low speed spool and the high speed spool occurs based on activation of the spool coupling system and reaching an engagement condition of the spool coupling system.Type: GrantFiled: August 9, 2019Date of Patent: October 4, 2022Assignee: RAYTHEON TECHNOLOGIES CORPORATIONInventors: Matthew R. Feulner, Gary Collopy
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Patent number: 11391205Abstract: A gas turbine engine is defined wherein the inlet guide vanes leading into a core engine flow path are sized and positioned such that flow paths positioned circumferentially intermediate the vane are sufficiently large that a hydraulic diameter of greater than or equal to about 1.5 is achieved. This will likely reduce the detrimental effect of icing.Type: GrantFiled: December 4, 2015Date of Patent: July 19, 2022Assignee: RAYTHEON TECHNOLOGIES CORPORATIONInventors: Matthew R. Feulner, Shengfang Liao
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Patent number: 11339678Abstract: A system for controlling blade tip clearances in a gas turbine engine may comprise an active clearance control system and a controller in operable communication with the active clearance control system. The controller may be configured to identify a cruise condition, reduce a thrust limit of the gas turbine engine to a de-rated maximum climb thrust, determine a first target tip clearance based on the de-rated maximum climb thrust, and send a command signal correlating to the first target tip clearance to the active clearance control system.Type: GrantFiled: July 19, 2018Date of Patent: May 24, 2022Assignee: Raytheon Technologies CorporationInventor: Matthew R. Feulner
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Publication number: 20210040896Abstract: A system includes a gas turbine engine having a low speed spool and a high speed spool. The system also includes a spool coupling system configured to mechanically link the low speed spool and the high speed spool. A controller is operable to determine a mode of operation of the gas turbine engine, monitor for a spool coupling activation condition associated with the mode of operation, and activate the spool coupling system based on the controller detecting the spool coupling activation condition. Engagement and power transfer between the low speed spool and the high speed spool occurs based on activation of the spool coupling system and reaching an engagement condition of the spool coupling system.Type: ApplicationFiled: August 9, 2019Publication date: February 11, 2021Inventors: Matthew R. Feulner, Gary Collopy
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Publication number: 20200332726Abstract: A method of operating a gas turbine engine includes commanding an acceleration of the gas turbine engine and moving a variable pitch high pressure compressor vane toward an open position thereby reducing an acceleration rate of a high pressure turbine rotor thereby reducing a change in a clearance gap between the high pressure turbine rotor and a blade outer airseal. An active clearance control system of a gas turbine engine includes an engine control system configured to command an acceleration of the gas turbine engine and move a variable pitch high pressure compressor vane toward an open position thereby slowing an acceleration rate of a high pressure turbine rotor thereby reducing a change in a clearance gap between the high pressure turbine rotor and a blade outer airseal located radially outboard of the high pressure turbine rotor.Type: ApplicationFiled: April 18, 2019Publication date: October 22, 2020Inventors: Joseph Kehoe, Richard P. Meisner, Manuj Dhingra, Patrick D. Couture, Matthew R. Feulner, Brenda J. Lisitano, Christopher L. Ho
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Patent number: 10787277Abstract: A system for starting a gas turbine engine of an aircraft is provided. The system includes a pneumatic starter motor, a discrete starter valve switchable between an on-state and an off-state, and a controller operable to perform a starting sequence for the gas turbine engine. The starting sequence includes alternating on and off commands to an electromechanical device coupled to the discrete starter valve to achieve a partially open position of the discrete starter valve to control a flow from a starter air supply to the pneumatic starter motor to drive rotation of a starting spool of the gas turbine engine below an engine idle speed, where the controller modulates a duty cycle of the discrete starter valve via pulse width modulation.Type: GrantFiled: July 13, 2018Date of Patent: September 29, 2020Assignee: RAYTHEON TECHNOLOGIES CORPORATIONInventors: Matthew D. Teicholz, Jeffrey W. Sutliff, William H. Greene, Jr., Kenneth J. White, Matthew R. Feulner, John P. Virtue, Jr., Jorn A. Glahn, Philip D. Hoover, Victor M. Pinedo, Jason B. Solomonides
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Patent number: 10774788Abstract: A gas turbine engine includes an engine core includes at least one compressor, a combustor downstream of the compressor, and at least one turbine downstream of the combustor. A primary flowpath fluidly connects each of the compressor, the combustor, and the turbine. At least one particle extraction duct has an extraction duct inlet connected to the primary flowpath fore of the compressor and an extraction duct outlet connected to a bypass flowpath.Type: GrantFiled: June 28, 2016Date of Patent: September 15, 2020Assignee: Raytheon Technologies CorporationInventor: Matthew R. Feulner
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Patent number: 10718272Abstract: A gas turbine engine comprises a housing having an inlet leading to a fan rotor. A bypass door is mounted upstream of the inlet to the fan rotor, and is moveable away from a non-bypass position to a bypass position to selectively bypass boundary layer air vertically beneath the engine. An aircraft is also disclosed.Type: GrantFiled: April 18, 2017Date of Patent: July 21, 2020Assignee: United Technologies CorporationInventors: Wesley K. Lord, Matthew R Feulner, Gabriel L. Suciu, Jesse M. Chandler
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Patent number: 10625881Abstract: A system for controlling a start sequence of a gas turbine engine includes an electronic engine control system, a thermal model, memory, a model for determining a time period (tmotoring), and a controller. The thermal model synthesizes a heat state of the gas turbine engine. The memory records the current heat state at shutdown and a shutdown time of the gas turbine engine. The model for determining the time period is for motoring the gas turbine engine at a predetermined speed Ntarget that is less than a speed to start the gas turbine engine, where tmotoring is a function of the heat state recorded at engine shutdown and an elapsed time of an engine start request relative to a previous shutdown time. The controller modulates a starter valve to maintain the gas turbine engine within a predetermined speed range of NtargetMin to NtargetMax for homogenizing engine temperatures.Type: GrantFiled: May 22, 2018Date of Patent: April 21, 2020Assignee: UNITED TECHNOLOGIES CORPORATIONInventors: Matthew D. Teicholz, Jeffrey W. Sutliff, William H. Greene, Jr., Kenneth J. White, Matthew R. Feulner, John P. Virtue, Jr., Jorn A. Glahn, Philip D. Hoover, Victor M. Pinedo, Jason B. Solomonides