Patents by Inventor Shankar S. Magge
Shankar S. Magge 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: 20170298832Abstract: A turbofan engine has an engine case and a gaspath through the engine case. A fan has a circumferential array of fan blades. The engine further has a compressor, a combustor, a gas generating turbine, and a low pressure turbine section. A speed reduction mechanism couples the low pressure turbine section to the fan. A bypass area ratio is greater than about 6.0. The low pressure turbine section airfoil count to bypass area ratio is below about 170.Type: ApplicationFiled: October 13, 2016Publication date: October 19, 2017Inventors: Paul R. Adams, Shankar S. Magge, Joseph Brent Staubach, Wesley K. Lord, Frederick M. Schwarz, Gabriel L. Suciu
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Publication number: 20170044978Abstract: A turbofan engine has an engine case and a gaspath through the engine case. A fan has a circumferential array of fan blades. The engine further has a compressor, a combustor, a gas generating turbine, and a low pressure turbine section. A speed reduction mechanism couples the low pressure turbine section to the fan. A bypass area ratio is greater than about 6.0. The low pressure turbine section airfoil count to bypass area ratio is below about 170.Type: ApplicationFiled: October 13, 2016Publication date: February 16, 2017Inventors: Paul R. Adams, Shankar S. Magge, Joseph Brent Staubach, Wesley K. Lord, Frederick M. Schwarz, Gabriel L. Suciu
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Publication number: 20170044992Abstract: A turbofan engine has an engine case and a gaspath through the engine case. A fan has a circumferential array of fan blades. The engine further has a compressor, a combustor, a gas generating turbine, and a low pressure turbine section. A speed reduction mechanism couples the low pressure turbine section to the fan. A bypass area ratio is greater than about 6.0. The low pressure turbine section airfoil count to bypass area ratio is below about 170.Type: ApplicationFiled: October 13, 2016Publication date: February 16, 2017Inventors: Paul R. Adams, Shankar S. Magge, Joseph Brent Staubach, Wesley K. Lord, Frederick M. Schwarz, Gabriel L. Suciu
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Publication number: 20170044990Abstract: A turbofan engine has an engine case and a gaspath through the engine case. A fan has a circumferential array of fan blades. The engine further has a compressor, a combustor, a gas generating turbine, and a low pressure turbine section. A speed reduction mechanism couples the low pressure turbine section to the fan. A bypass area ratio is greater than about 6.0. The low pressure turbine section airfoil count to bypass area ratio is below about 170.Type: ApplicationFiled: October 13, 2016Publication date: February 16, 2017Inventors: Paul R. Adams, Shankar S. Magge, Joseph Brent Staubach, Wesley K. Lord, Frederick M. Schwarz, Gabriel L. Suciu
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Publication number: 20160222815Abstract: A gas turbine engine comprises a fan drive turbine. The fan drive turbine drives the fan through a gear reduction; A change in enthalpy is defined across the gas turbine engine. The change in enthalpy divided by a speed of the fan drive turbine squared is less than or equal to about 1.8. An axial component of gases approaching an upstream most blade of the fan drive turbine divided by the speed of the fan drive turbine is equal to or less than about 0.9.Type: ApplicationFiled: March 5, 2014Publication date: August 4, 2016Inventors: Frederick M. Schwarz, Shankar S. Magge
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Publication number: 20150377122Abstract: A turbofan engine includes an engine case, a gaspath through the engine case, a fan having an array of fan blades, a compressor in fluid communication with the fan, a combustor in fluid communication with the compressor, and a turbine in fluid communication with the combustor. The turbine has a fan drive turbine section having 3 to 6 blade stages. A speed reduction mechanism couples the fan drive turbine section to the fan. A ratio of maximum gaspath radius along the low pressure turbine section to maximum radius of the fan blades is less than about 0.55. A bypass area ratio is greater than about 6.0. A ratio of a fan drive turbine section airfoil count to the bypass area ratio is less than about 170 and a second turbine section.Type: ApplicationFiled: July 8, 2015Publication date: December 31, 2015Inventors: Paul R. Adams, Shankar S. Magge, Joseph Brent Staubach, Wesley K. Lord, Frederick M. Schwarz, Gabriel L. Suciu
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Publication number: 20150377124Abstract: A turbofan engine includes an engine case, a gaspath through the engine case, a fan having a circumferential array of fan blades, a compressor in fluid communication with the fan, a combustor in fluid communication with the compressor, and a turbine in fluid communication with the combustor. The turbine has a fan drive turbine section having 3 to 6 blade stages. A speed reduction mechanism couples the fan drive turbine section to the fan. A bypass area ratio is between about 8.0 and about 20.0. A ratio of maximum gaspath radius along the fan drive turbine section to maximum radius of the fan is less than about 0.50. A ratio of a turbine section airfoil count to the bypass area ratio is between about 10 and about 170. The fan drive turbine section airfoil count being the total number of blade airfoils and vane airfoils of the fan drive turbine section.Type: ApplicationFiled: July 8, 2015Publication date: December 31, 2015Inventors: Paul R. Adams, Shankar S. Magge, Joseph Brent Staubach, Wesley K. Lord, Frederick M. Schwarz, Gabriel L. Suciu
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Publication number: 20150377123Abstract: A turbofan engine comprises a fan having fan blades. A compressor is in communication with the fan section. The fan is configured to communicate a portion of air into a bypass path defining a bypass area outwardly of the compressor and a portion into the compressor. A bypass ratio is defined as air communicated through the bypass path relative to air communicated to the compressor being greater than about 6.0. A combustor is in fluid communication with the compressor. A turbine is in communication with the combustor. The turbine has a first turbine section that includes two or more stages and a second turbine section that includes at least two stages. A ratio of airfoils in the first turbine section to the bypass ratio is less than about 170. The first turbine section includes a maximum gas path radius. A ratio of the maximum gas path radius to a maximum radius of the fan blades is less than about 0.50. A speed reduction mechanism is coupled to the fan and rotatable by the turbine.Type: ApplicationFiled: July 8, 2015Publication date: December 31, 2015Inventors: Paul R. Adams, Shankar S. Magge, Joseph Brent Staubach, Wesley K. Lord, Frederick M. Schwarz, Gabriel L. Suciu
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Publication number: 20150345404Abstract: A turbofan engine has an engine case and a gaspath through the engine case. A fan has a circumferential array of fan blades. The engine further has a compressor, a combustor, a gas generating turbine, and a low pressure turbine section. A speed reduction mechanism couples the low pressure turbine section to the fan. A bypass area ratio is greater than about 6.0. The low pressure turbine section airfoil count to bypass area ratio is below about 170.Type: ApplicationFiled: April 21, 2015Publication date: December 3, 2015Inventors: Paul R. Adams, Shankar S. Magge, Joseph B. Staubach, Wesley K. Lord, Frederick M. Schwarz
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Patent number: 9028200Abstract: A gas turbine engine includes a shaft defining an axis of rotation. An outer turbine rotor directly drives the shaft and includes an outer set of blades. An inner turbine rotor has an inner set of blades interspersed with the outer set of blades. The inner turbine rotor is configured to rotate in an opposite direction about the axis of rotation from the outer turbine rotor. A splitter gear system couples the inner turbine rotor to the shaft and is configured to rotate the inner set of blades at a faster speed than the outer set of blades.Type: GrantFiled: February 29, 2012Date of Patent: May 12, 2015Assignee: United Technologies CorporationInventors: Gabriel L. Suciu, Brian D. Merry, Michael E. McCune, Shankar S. Magge
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Patent number: 9010085Abstract: A turbofan engine has an engine case and a gaspath through the engine case. A fan has a circumferential array of fan blades. The engine further has a compressor, a combustor, a gas generating turbine, and a low pressure turbine section. A speed reduction mechanism couples the low pressure turbine section to the fan. A bypass area ratio is greater than about 6.0. The low pressure turbine section airfoil count to bypass area ratio ratio is below about 170.Type: GrantFiled: August 30, 2012Date of Patent: April 21, 2015Assignee: United Technologies CorporationInventors: Paul R. Adams, Shankar S. Magge, Joseph B. Staubach, Wesley K. Lord, Frederick M. Schwarz, Gabriel L. Suciu
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Patent number: 8935926Abstract: An impeller includes a plurality of vanes formed around a hub, each of the plurality of vanes defines an offset between a leading edge and a trailing edge.Type: GrantFiled: October 28, 2010Date of Patent: January 20, 2015Assignee: United Technologies CorporationInventors: Joel H. Wagner, Shankar S. Magge, Keith A. Santeler
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Patent number: 8915090Abstract: A gas turbine engine includes first and second stages having a rotational axis. A circumferential array of airfoils is arranged axially between the first stage and the second stage. At least one of the airfoils have a curvature provided equidistantly between pressure and suction sides. The airfoils extend from a leading edge to a trailing edge at a midspan plane along the airfoil. An angle is defined between first and second lines respectively tangent to the intersection of the midspan plane and the curvature at airfoil leading and trailing edges. The angle is equal to or greater than about 10°.Type: GrantFiled: March 21, 2014Date of Patent: December 23, 2014Assignee: United Technologies CorporationInventors: Thomas J. Praisner, Shankar S. Magge, Matthew B. Estes
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Patent number: 8850793Abstract: A turbofan engine is disclosed and includes a fan and a compressor in communication with the fan section, a combustor, a turbine and a speed reduction mechanism coupled to the fan and rotatable by the turbine. The turbine includes a first turbine section that includes three or more stages and a second turbine section that includes at least two stages. A ratio of airfoils in the first turbine section to a bypass area is less than about 170.Type: GrantFiled: December 11, 2013Date of Patent: October 7, 2014Assignee: United Technologies CorporationInventors: Paul R. Adams, Shankar S. Magge, Joseph B. Staubach, Wesley K. Lord, Frederick M. Schwarz
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Patent number: 8844265Abstract: A turbofan engine has an engine case and a gaspath through the engine case. A fan has a circumferential array of fan blades. The engine further has a compressor, a combustor, a gas generating turbine, and a low pressure turbine section. A speed reduction mechanism couples the low pressure turbine section to the fan. A bypass area ratio is greater than about 6.0. The low pressure turbine section airfoil count to bypass area ratio is below about 170.Type: GrantFiled: May 18, 2012Date of Patent: September 30, 2014Assignee: United Technologies CorporationInventors: Paul R. Adams, Shankar S. Magge, Joseph B. Staubach, Wesley K. Lord, Frederick M. Schwarz, Gabriel L. Suciu
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Publication number: 20140202133Abstract: A gas turbine engine includes first and second stages having a rotational axis. A circumferential array of airfoils is arranged axially between the first stage and the second stage. At least one of the airfoils have a curvature provided equidistantly between pressure and suction sides. The airfoils extend from a leading edge to a trailing edge at a midspan plane along the airfoil. An angle is defined between first and second lines respectively tangent to the intersection of the midspan plane and the curvature at airfoil leading and trailing edges. The angle is equal to or greater than about 10°.Type: ApplicationFiled: March 21, 2014Publication date: July 24, 2014Applicant: United Technologies CorporationInventors: Thomas J. Praisner, Shankar S. Magge, Matthew B. Estes
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Publication number: 20140174055Abstract: A turbofan engine is disclosed and includes a fan and a compressor in communication with the fan section, a combustor, a turbine and a speed reduction mechanism coupled to the fan and rotatable by the turbine. The turbine includes a first turbine section that includes three or more stages and a second turbine section that includes at least two stages. A ratio of airfoils in the first turbine section to a bypass area is less than about 170.Type: ApplicationFiled: December 11, 2013Publication date: June 26, 2014Applicant: United Technologies CorporationInventors: Paul R. Adams, Shankar S. Magge, Joseph B. Staubach, Wesley K. Lord, Frederick M. Schwarz
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Publication number: 20140102076Abstract: A turbofan engine has an engine case and a gaspath through the engine case. A fan has a circumferential array of fan blades. The engine further has a compressor, a combustor, a gas generating turbine, and a low pressure turbine section. A speed reduction mechanism couples the low pressure turbine section to the fan. A bypass area ratio is greater than about 6.0. The low pressure turbine section airfoil count to bypass area ratio ratio is below about 170.Type: ApplicationFiled: August 30, 2012Publication date: April 17, 2014Applicant: UNITED TECHNOLOGIES CORPORATIONInventors: Paul R. Adams, Shankar S. Magge, Joseph B. Staubach, Wesley K. Lord, Frederick M. Schwarz, Gabriel L. Suciu
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Patent number: 8561414Abstract: A gas turbine engine includes first and second stages having a rotational axis. A circumferential array of airfoils is arranged axially between the first stage and the second stage. At least one of the airfoils have a curvature provided equidistantly between pressure and suction sides. The airfoils extend from a leading edge to a trailing edge at a midspan plane along the airfoil. An angle is defined between first and second lines respectively tangent to the intersection of the midspan plane and the curvature at airfoil leading and trailing edges. The angle is equal to or greater than about 10°.Type: GrantFiled: February 27, 2013Date of Patent: October 22, 2013Assignee: United Technologies CorporationInventors: Thomas J. Praisner, Shankar S. Magge, Matthew B. Estes
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Publication number: 20130223983Abstract: A gas turbine engine includes a shaft defining an axis of rotation. An outer turbine rotor directly drives the shaft and includes an outer set of blades. An inner turbine rotor has an inner set of blades interspersed with the outer set of blades. The inner turbine rotor is configured to rotate in an opposite direction about the axis of rotation from the outer turbine rotor. A splitter gear system couples the inner turbine rotor to the shaft and is configured to rotate the inner set of blades at a faster speed than the outer set of blades.Type: ApplicationFiled: February 29, 2012Publication date: August 29, 2013Inventors: Gabriel L. Suciu, Brian D. Merry, Michael E. McCune, Shankar S. Magge