Patents by Inventor Bret M. Teller
Bret M. Teller 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: 20190292940Abstract: A turbine vane for a gas turbine engine having a plurality of cooling holes defined therein, the plurality of cooling holes provide fluid communication to a surface of the turbine vane, the plurality of cooling holes including holes noted by the following coordinates: HDA, HDB, HEA, HEB, SAA, SAB, and HCA of Table 1.Type: ApplicationFiled: March 23, 2018Publication date: September 26, 2019Inventors: Christopher Cosher, Alex J. Schneider, Bret M. Teller
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Patent number: 10406596Abstract: A casting core for an airfoil according to an example of the present disclosure includes, among other things, a first portion extending in a first direction and corresponding to a first cavity of an airfoil. The first portion defines a reference plane along a parting line formed by a casting die. The first portion defines a plurality of grooves corresponding to a plurality of trip strips of the airfoil. Each of the plurality of grooves defines a respective groove axis, and the plurality of grooves are distributed in the first direction along a first side of the reference plane such that one or more of the groove axes are oriented with respect to a pull direction of the casting die. A method for fabricating a gas turbine engine component is also disclosed.Type: GrantFiled: May 1, 2015Date of Patent: September 10, 2019Assignee: UNITED TECHNOLOGIES CORPORATIONInventors: Matthew S. Gleiner, Bret M. Teller, James T. Auxier
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Patent number: 10408071Abstract: A vane according to an exemplary aspect of the present disclosure includes, among other things, a platform extending from an edge face and between spaced apart lateral faces and an airfoil extending outwardly from the platform. The platform includes at least one ejection port in the edge face and at least one passage connected to the at least one ejection port.Type: GrantFiled: August 26, 2014Date of Patent: September 10, 2019Assignee: UNITED TECHNOLOGIES CORPORATIONInventors: Scott D. Lewis, Dominic J. Mongillo, Jr., Mark F. Zelesky, Bret M. Teller, Russell Deibel, Matthew S. Gleiner
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Patent number: 10364680Abstract: A component for a gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, a platform that axially extends between a leading edge and a trailing edge and circumferentially extends between a first mate face and a second mate face and a trench disposed on at least one of the first mate face and the second mate face. A plurality of cooling holes are axially disposed within the trench.Type: GrantFiled: August 14, 2012Date of Patent: July 30, 2019Assignee: UNITED TECHNOLOGIES CORPORATIONInventors: Bret M. Teller, Mark F. Zelesky, Scott D. Lewis, Brandon W. Spangler, Ricardo Trindade
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Patent number: 10337332Abstract: An airfoil of a gas turbine engine includes an airfoil body having a leading edge and a trailing edge extending in a radial direction, a trailing edge cavity formed within the airfoil and proximate to the trailing edge of the airfoil, the trailing edge cavity extending from the trailing edge in a forward direction toward the leading edge, at least one set of blocking pedestals located within the trailing edge cavity, a set of circular pedestals located aftward from the at least one blocking set of pedestals, and a set of spear pedestals located aftward from the set of circular pedestals and closest to the trailing edge of the airfoil body.Type: GrantFiled: February 25, 2016Date of Patent: July 2, 2019Assignee: UNITED TECHNOLOGIES CORPORATIONInventors: James Tilsley Auxier, Bret M. Teller
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Patent number: 10227875Abstract: A component for a gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, a platform that axially extends between a leading edge and a trailing edge, circumferentially extends between a first mate face and a second mate face, and includes a gas path surface and a non-gas path surface. The component defines at least one cavity that extends at least partially inside of the component. A first plurality of cooling holes extends from the at least one cavity to at least one of the first mate face and the second mate face and a second plurality of cooling holes extends from either the at least one cavity or the non-gas path surface to the gas path surface.Type: GrantFiled: February 7, 2014Date of Patent: March 12, 2019Assignee: UNITED TECHNOLOGIES CORPORATIONInventors: Scott D. Lewis, Brandon M. Rapp, Jeffrey S. Beattie, Matthew Andrew Hough, Bret M. Teller, Jeffrey Michael Jacques, Max Asterlin
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Patent number: 10180067Abstract: A gas turbine engine component comprises a shroud, a U-channel, an internal cooling air passage and a U-channel cooling hole. The shroud comprises a forward face, an aft face, a first side face and a second side face. The U-channel is disposed in the aft face of the shroud. A gas path surface connects the forward face, aft face, first side face and second side face. A cooled surface connects the forward face, aft face, first side face and second side face opposite the gas path face. The internal cooling air passage extends through the shroud. The U-channel cooling hole extends into the first side face of the shroud adjacent the U-channel to intersect the internal cooling passage.Type: GrantFiled: May 31, 2012Date of Patent: January 15, 2019Assignee: United Technologies CorporationInventors: Jeffrey S. Beattie, Scott D. Lewis, Mark F. Zelesky, Ricardo Trindade, Bret M. Teller, Jeffrey Michael Jacques, Brandon M. Rapp
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Patent number: 9797260Abstract: A gas turbine engine component according to an exemplary aspect of this disclosure includes a peripheral portion defining an internal cooling passage and a recess in fluid communication with the internal cooling passage. The peripheral portion includes an outer wear surface, and the recess tapers toward the outer wear surface.Type: GrantFiled: April 23, 2015Date of Patent: October 24, 2017Assignee: UNITED TECHNOLOGIES CORPORATIONInventors: Graham Ryan Philbrick, Ken Lagueux, Bret M. Teller
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Publication number: 20170292384Abstract: A casting core and/or an airfoil may comprise a tip flag cavity having a forward pedestal and a first spear pedestal disposed aft of the forward pedestal. A trailing edge discharge cavity may be separated from the tip flag cavity and include a first row of pedestals. The first row of pedestals may comprise a first racetrack pedestal. A second row of pedestals may be disposed aft of the first row of pedestals and include a second racetrack pedestal. A third row of pedestals may be disposed aft of the second row of pedestals and include a circular pedestal. A fourth row of pedestals may be disposed aft of the third row of pedestals and include a second spear pedestal.Type: ApplicationFiled: April 11, 2016Publication date: October 12, 2017Applicant: United Technologies CorporationInventors: James T. Auxier, Parth Jariwala, Bret M. Teller
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Publication number: 20170248021Abstract: An airfoil of a gas turbine engine includes an airfoil body having a leading edge and a trailing edge extending in a radial direction, a trailing edge cavity formed within the airfoil and proximate to the trailing edge of the airfoil, the trailing edge cavity extending from the trailing edge in a forward direction toward the leading edge, at least one set of blocking pedestals located within the trailing edge cavity, a set of circular pedestals located aftward from the at least one blocking set of pedestals, and a set of spear pedestals located aftward from the set of circular pedestals and closest to the trailing edge of the airfoil body.Type: ApplicationFiled: February 25, 2016Publication date: August 31, 2017Inventors: James Tilsley Auxier, Bret M. Teller
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Publication number: 20170211396Abstract: An airfoil of a gas turbine engine is provided including a leading edge extending in a radial direction, a tip extending in an axial direction from the leading edge, a first rib extending radially within the airfoil, the leading edge and the first rib defining a leading edge cavity within the airfoil, a second rib, the second rib and the first rib defining a serpentine cavity therein, a third rib extending axially within the tip, a flag tip cavity defined by the third rib, the leading edge, and the tip, the leading edge cavity fluidly connected to the flag tip cavity, and a bypass aperture formed between the first rib and the third rib, the bypass aperture configured to fluidly connect the serpentine cavity with the flag tip cavity.Type: ApplicationFiled: May 2, 2016Publication date: July 27, 2017Inventors: Dominic J. Mongillo, Parth Jariwala, Bret M. Teller, Mark F. Zelesky, James Tilsley Auxier
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Publication number: 20160319674Abstract: A casting core for an airfoil according to an example of the present disclosure includes, among other things, a first portion extending in a first direction and corresponding to a first cavity of an airfoil. The first portion defines a reference plane along a parting line formed by a casting die. The first portion defines a plurality of grooves corresponding to a plurality of trip strips of the airfoil. Each of the plurality of grooves defines a respective groove axis, and the plurality of grooves are distributed in the first direction along a first side of the reference plane such that one or more of the groove axes are oriented with respect to a pull direction of the casting die. A method for fabricating a gas turbine engine component is also disclosed.Type: ApplicationFiled: May 1, 2015Publication date: November 3, 2016Inventors: Matthew S. Gleiner, Bret M. Teller, James T. Auxier
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Publication number: 20160230577Abstract: A vane according to an exemplary aspect of the present disclosure includes, among other things, a platform extending from an edge face and between spaced apart lateral faces and an airfoil extending outwardly from the platform. The platform includes at least one ejection port in the edge face and at least one passage connected to the at least one ejection port.Type: ApplicationFiled: August 26, 2014Publication date: August 11, 2016Applicant: United Technologies CorporationInventors: Scott D. Lewis, Dominic J. Mongillo, JR., Mark F. Zelesky, Bret M. Teller, Russell Deibel, Matthew S. Gleiner
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Publication number: 20160090844Abstract: A component for a gas turbine engine includes an exterior surface that provides pressure and suction sides. A cooling passage in the component includes a serpentine passageway that has first and second passes respectively configured to provide fluid flow in opposite directions from one another. The first pass includes first and second portions nested relative to one another and overlapping in a thickness direction. The first and second portions are adjacent to one another by sharing a common wall. The first portion is provided on the suction side. The second portion is provided on the pressure side.Type: ApplicationFiled: April 15, 2015Publication date: March 31, 2016Inventors: James T. Auxier, Parth Jariwala, Mark F. Zelesky, Bret M. Teller
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Patent number: 9243500Abstract: A turbine blade for a gas turbine engine includes an airfoil including leading and trailing edges joined by spaced apart pressure and suction sides to provide an exterior airfoil surface extending in a radial direction. The trailing edge is arranged on an aft side of the turbine blade. A root supports a platform from which the airfoil extends and a cooling passage extends within the root in the radial direction to the airfoil. A lower wing is arranged beneath the platform on the aft side and extends in an axial direction to provide a U-shaped channel with the platform that extends in a circumferential direction. An impingement hole extends from the U-channel to the cooling passage.Type: GrantFiled: June 29, 2012Date of Patent: January 26, 2016Assignee: UNITED TECHNOLOGIES CORPORATIONInventors: Mark F. Zelesky, Ricardo Trindade, Bret M. Teller, Scott D. Lewis, Jeffrey S. Beattie, Jeffrey Michael Jacques, Brandon M. Rapp
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Publication number: 20150377032Abstract: A component for a gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, a platform that axially extends between a leading edge and a trailing edge, circumferentially extends between a first mate face and a second mate face, and includes a gas path surface and a non-gas path surface. The component defines at least one cavity that extends at least partially inside of the component. A first plurality of cooling holes extends from the at least one cavity to at least one of the first mate face and the second mate face and a second plurality of cooling holes extends from either the at least one cavity or the non-gas path surface to the gas path surface.Type: ApplicationFiled: February 7, 2014Publication date: December 31, 2015Inventors: Scott D. LEWIS, Brandon M. RAPP, Jeffrey S. BEATTIE, Matthew Andrew HOUGH, Bret M. TELLER, Jeffrey Michael JACQUES, Max ASTERLIN
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Publication number: 20150308274Abstract: A gas turbine engine component according to an exemplary aspect of this disclosure includes a peripheral portion defining an internal cooling passage and a recess in fluid communication with the internal cooling passage. The peripheral portion includes an outer wear surface, and the recess tapers toward the outer wear surface.Type: ApplicationFiled: April 23, 2015Publication date: October 29, 2015Inventors: Graham Ryan Philbrick, Ken Lagueux, Bret M. Teller
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Publication number: 20140321961Abstract: A gas turbine engine component comprises a shroud, a U-channel, an internal cooling air passage and a U-channel cooling hole. The shroud comprises a forward face, an aft face, a first side face and a second side face. The U-channel is disposed in the aft face of the shroud. A gas path surface connects the forward face, aft face, first side face and second side face. A cooled surface connects the forward face, aft face, first side face and second side face opposite the gas path face. The internal cooling air passage extends through the shroud. The U-channel cooling hole extends into the first side face of the shroud adjacent the U-channel to intersect the internal cooling passage.Type: ApplicationFiled: May 31, 2012Publication date: October 30, 2014Applicant: UNITED TECHNOLOGIES CORPORATIONInventors: Jeffrey S. Beattie, Scott D. Lewis, Mark F. Zelesky, Ricardo Trindade, Bret M. Teller, Jeffrey Michael Jacques, Brandon M. Rapp
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Publication number: 20140047844Abstract: A component for a gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, a platform that axially extends between a leading edge and a trailing edge and circumferentially extends between a first mate face and a second mate face and a trench disposed on at least one of the first mate face and the second mate face. A plurality of cooling holes are axially disposed within the trench.Type: ApplicationFiled: August 14, 2012Publication date: February 20, 2014Inventors: Bret M. Teller, Mark F. Zelesky, Scott D. Lewis, Brandon W. Spangler, Ricardo Trindade
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Publication number: 20140000282Abstract: A turbine blade for a gas turbine engine includes an airfoil including leading and trailing edges joined by spaced apart pressure and suction sides to provide an exterior airfoil surface extending in a radial direction. The trailing edge is arranged on an aft side of the turbine blade. A root supports a platform from which the airfoil extends and a cooling passage extends within the root in the radial direction to the airfoil. A lower wing is arranged beneath the platform on the aft side and extends in an axial direction to provide a U-shaped channel with the platform that extends in a circumferential direction. An impingement hole extends from the U-channel to the cooling passage.Type: ApplicationFiled: June 29, 2012Publication date: January 2, 2014Inventors: Mark F. Zelesky, Ricardo Trindade, Bret M. Teller, Scott D. Lewis, Jeffrey S. Beattie, Jeffrey Michael Jacques, Brandon M. Rapp