Patents by Inventor Gary Roberge
Gary Roberge 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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Patent number: 11530635Abstract: A fluid injection system for a gas turbine engine may comprise a fluid injector configured to inject a fluid into an exhaust flow exiting a turbine section of the gas turbine engine. The fluid injector may be coupled to a turbine exit guide vane located at a forward end of an exhaust system of the gas turbine engine. The fluid may decrease a temperature of the exhaust flow exiting the turbine section and/or increase a thrust of the gas turbine engine.Type: GrantFiled: September 27, 2019Date of Patent: December 20, 2022Assignee: Raytheon Technologies CorporationInventors: Daniel B. Kupratis, Neil Terwilliger, Gary Roberge
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Patent number: 11377219Abstract: A hybrid electric gas turbine propulsion system may comprise: a first propulsion system, a second propulsion system, and a third propulsion system. The first propulsion system may comprise a first fan, a first turbine, a first compressor, and a first electric motor, the first fan operably coupled to the first turbine and the first compressor by a first shaft, the first shaft coupled to the first electric motor, the first shaft configured to be disposed radially inward of a fuselage of an aircraft. The second propulsion system and the third propulsion system may be in accordance with the first propulsion system. The hybrid electric gas turbine propulsion system may be symmetric about a vertical plane extending through a neutral aerodynamic axis.Type: GrantFiled: April 17, 2020Date of Patent: July 5, 2022Assignee: Raytheon Technologies CorporationInventors: Daniel Kupratis, Gary Roberge
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Publication number: 20210323685Abstract: A hybrid electric gas turbine propulsion system may comprise: a first propulsion system, a second propulsion system, and a third propulsion system. The first propulsion system may comprise a first fan, a first turbine, a first compressor, and a first electric motor, the first fan operably coupled to the first turbine and the first compressor by a first shaft, the first shaft coupled to the first electric motor, the first shaft configured to be disposed radially inward of a fuselage of an aircraft. The second propulsion system and the third propulsion system may be in accordance with the first propulsion system. The hybrid electric gas turbine propulsion system may be symmetric about a vertical plane extending through a neutral aerodynamic axis.Type: ApplicationFiled: April 17, 2020Publication date: October 21, 2021Applicant: RAYTHEON TECHNOLOGIES CORPORATIONInventors: Daniel Kupratis, Gary Roberge
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Publication number: 20200182113Abstract: A fluid injection system for a gas turbine engine may comprise a fluid injector configured to inject a fluid into an exhaust flow exiting a turbine section of the gas turbine engine. The fluid injector may be coupled to a turbine exit guide vane located at a forward end of an exhaust system of the gas turbine engine. The fluid may decrease a temperature of the exhaust flow exiting the turbine section and/or increase a thrust of the gas turbine engine.Type: ApplicationFiled: September 27, 2019Publication date: June 11, 2020Applicant: UNITED TECHNOLOGIES CORPORATIONInventors: Daniel B. Kupratis, Neil Terwilliger, Gary Roberge
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Publication number: 20190048727Abstract: A gas turbine engine component includes an airfoil that has first and second structural airfoil segments that are bonded to each other in at least one diffusion joint. The first and second structural airfoil segments are formed of, respectively, first and second materials. The first and second materials are: different base-metal metallic alloys, a metallic alloy and a ceramic-based material, or ceramic-based materials that differ by at least one of composition and microstructure. The first structural airfoil segment is a first skin and the second structural airfoil segment is a hollow core that has an airfoil shape.Type: ApplicationFiled: October 16, 2018Publication date: February 14, 2019Inventors: Gary Roberge, Grant O. Cook, III
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Patent number: 10145245Abstract: A component includes a component body that is configured for use in a gas turbine engine. The component body includes first and second structural segments that are bonded to each other in at least one diffusion joint. The first and second structural segments are formed of, respectively, first and second materials. The first and second materials are different base-metal alloys, a metallic alloy and a ceramic-based material, or ceramic-based materials that differ by at least one of composition and microstructure.Type: GrantFiled: September 8, 2014Date of Patent: December 4, 2018Assignee: United Technologies CorporationInventors: Gary Roberge, Grant O. Cook, III
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Patent number: 9909494Abstract: A tip turbine engine includes an axial compressor having a plurality of airfoils compressing core airflow. The airfoils include bleed air openings on their suction side surfaces. The bleed air openings prevent separation of the compressed airflow, which permits each airfoil stage to perform increased compression without separation of the airflow. As a result, the number of stages can be reduced, thereby shortening the overall length of the turbine engine.Type: GrantFiled: February 15, 2006Date of Patent: March 6, 2018Assignee: United Technologies CorporationInventor: Gary Roberge
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Publication number: 20160215627Abstract: A component includes a component body that is configured for use in a gas turbine engine. The component body includes first and second structural segments that are bonded to each other in at least one diffusion joint. The first and second structural segments are formed of, respectively, first and second materials. The first and second materials are different base-metal alloys, a metallic alloy and a ceramic-based material, or ceramic-based materials that differ by at least one of composition and microstructure.Type: ApplicationFiled: September 8, 2014Publication date: July 28, 2016Inventors: Gary Roberge, Grant O. Cook, III
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Patent number: 8104257Abstract: A tip turbine engine (10) provides first and second turbines (32) rotatably driven by a combustor (30) generating a high-energy gas stream. The first turbine (32) is mounted at an outer periphery of a first fan (24a), such that the first fan is rotatably driven by the first turbine (32a). The second turbine (32b) is mounted at an outer periphery of a second fan (24b), and is rotatably driven by the high-energy gas stream. In one embodiment, the first turbine (32a) rotatably drives a plurality of stages of first compressor blades (54) in an axial compressor (22) in a first rotational direction, while the second turbine (32b) rotatably drives a plurality of stages of second compressor blades (52) in the axial compressor (22) in a second rotational direction opposite the first. By rotatably driving alternating stages of compressor blades in opposite directions, the efficiency of the axial compressor (22) is increased and/or the number of stages of compressor blades can be reduced.Type: GrantFiled: December 1, 2004Date of Patent: January 31, 2012Assignee: United Technologies CorporationInventors: James W. Norris, Craig A. Nordeen, Gary Roberge
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Publication number: 20090145136Abstract: A tip turbine engine (10) provides first and second turbines (32) rotatably driven by a combustor (30) generating a high-energy gas stream. The first turbine (32) is mounted at an outer periphery of a first fan (24a), such that the first fan is rotatably driven by the first turbine (32a). The second turbine (32b) is mounted at an outer periphery of a second fan (24b), and is rotatably driven by the high-energy gas stream. In one embodiment, the first turbine (32a) rotatably drives a plurality of stages of first compressor blades (54) in an axial compressor (22) in a first rotational direction, while the second turbine (32b) rotatably drives a plurality of stages of second compressor blades (52) in the axial compressor (22) in a second rotational direction opposite the first. By rotatably driving alternating stages of compressor blades in opposite directions, the efficiency of the axial compressor (22) is increased and/or the number of stages of compressor blades can be reduced.Type: ApplicationFiled: December 1, 2004Publication date: June 11, 2009Inventors: James W. Norris, Craig A. Nordeen, Gary Roberge
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Publication number: 20090019858Abstract: A tip turbine engine includes an axial compressor having a plurality of airfoils compressing core airflow. The airfoils include bleed air openings on their suction side surfaces. The bleed air openings prevent separation of the compressed airflow, which permits each airfoil stage to perform increased compression without separation of the airflow. As a result, the number of stages can be reduced, thereby shortening the overall length of the turbine engine.Type: ApplicationFiled: February 15, 2006Publication date: January 22, 2009Inventor: Gary Roberge
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Publication number: 20080014078Abstract: A tip turbine engine (10) includes a combustor (30) radially outward of a fan. In order to reduce the heat transfer from the combustor and the high-energy gas stream generated by the combustor, a cold air ejector (38) radially outward of the combustor extends from a forward end of the nacelle (12) to a point rearward of the combustor and an exhaust mixer (110). The cold air ejector includes an annular inlet (17) at the forward end of the nacelle. The cold air ejector draws air over the outer engine case (39) to provide a boundary between the nacelle and the hot outer engine case. The layer of air being pulled past the engine case ejects the heat, thereby preventing the heat from escaping into the nacelle or engine bay.Type: ApplicationFiled: December 1, 2004Publication date: January 17, 2008Inventors: Gabriel Suciu, Gary Roberge
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Publication number: 20070144141Abstract: A propulsion system for producing thrust includes a turbine engine. The engine has a compressor, a combustor downstream of the compressor along a flow path, and a turbine downstream of the combustor along the flow path. A plurality of combustion conduits have outlets positioned to discharge combustion gas to form a fluidic nozzle throat.Type: ApplicationFiled: December 22, 2005Publication date: June 28, 2007Inventors: Gary Roberge, James Wozniak, Wendell Twelves
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Patent number: 7201558Abstract: A fan-turbine rotor hub includes an outer periphery scalloped by a multitude of elongated openings. Each elongated opening defines an inducer receipt section to receive an inducer section and a hollow fan blade section. An inducer exit from each inducer section is located adjacent a core airflow passage within each fan blade section to provide communication therebetween. A seal is located between an inner fan blade mount and a blade receipt section to minimize airflow leakage between the inducer exit and the core airflow passage.Type: GrantFiled: May 5, 2005Date of Patent: April 10, 2007Assignee: United Technologies CorporationInventors: James W. Norris, Craig A. Nordeen, Gary Roberge, Gabriel Suciu
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Publication number: 20070022738Abstract: A fan-turbine rotor assembly includes a diffuser mountable to the outer periphery of a multitude of fan blade sections to provide structural support to the outer tips of the fan blade sections and to turn and diffuse the airflow from the radial core airflow passage toward an axial airflow direction. The diffuser includes a fan blade tip shroud inner portion which forms a planar ring about the multiple of fan blade sections. Diffuser support rings are mounted to the fan blade tip shroud to further share the radial load applied to the fan blade sections by the blade mounted diffuser and annular turbine.Type: ApplicationFiled: July 27, 2005Publication date: February 1, 2007Inventors: James Norris, Craig Nordeen, Gary Roberge, Brian Merry
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Publication number: 20060251508Abstract: A fan-turbine rotor hub includes an outer periphery scalloped by a multitude of elongated openings. Each elongated opening defines an inducer receipt section to receive an inducer section and a hollow fan blade section. An inducer exit from each inducer section is located adjacent a core airflow passage within each fan blade section to provide communication therebetween. A seal is located between an inner fan blade mount and a blade receipt section to minimize airflow leakage between the inducer exit and the core airflow passage.Type: ApplicationFiled: May 5, 2005Publication date: November 9, 2006Inventors: James Norris, Craig Nordeen, Gary Roberge, Gabriel Suciu