Patents by Inventor James Harvey Laflen
James Harvey Laflen 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: 8226360Abstract: A turbine nozzle includes a row of vanes extending radially between annular outer and inner bands. The outer band includes a pair of radial flanges defining an annular seal groove therebetween. One of the flanges is crenelated to improve nozzle life.Type: GrantFiled: October 31, 2008Date of Patent: July 24, 2012Assignee: General Electric CompanyInventors: Patrick Jarvis Scoggins, James Harvey Laflen, Ching-Pang Lee, Wilson Frost
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Patent number: 8147192Abstract: A turbine shroud includes a shroud hanger having an arcuate panel from which three inner hooks extend inwardly, and from which two outer hooks extend outwardly therefrom. The two outer hooks effect a statically determinate configuration of the shroud.Type: GrantFiled: September 19, 2008Date of Patent: April 3, 2012Assignee: General Electric CompanyInventors: Daniel Vern Jones, James Harvey Laflen, Richard William Albrecht, Dustin Alfred Placke, Ching-Pang Lee
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Patent number: 7785067Abstract: A method of assembling a gas turbine engine is provided. The method includes coupling at least one turbine nozzle segment within the gas turbine engine. The at least one turbine nozzle segment includes at least one airfoil vane extending between an inner band and an outer band that includes an aft flange and a radial inner surface. The method also includes coupling at least one turbine shroud segment downstream from the at least one turbine nozzle segment, wherein the at least one turbine shroud segment includes a leading edge and a radial inner surface, and coupling a cooling fluid source in flow communication with the at least one turbine nozzle segment such that cooling fluid channeled to each turbine nozzle outer band aft flange is directed at an oblique discharge angle towards the leading edge of the at least one turbine shroud segment.Type: GrantFiled: November 30, 2006Date of Patent: August 31, 2010Assignee: General Electric CompanyInventors: Ching-Pang Lee, Eric Alan Estill, James Harvey Laflen, Paul Hadley Vitt, Michael Elliot Wymore
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Patent number: 7740444Abstract: A method for cooling a turbine shroud assembly includes providing a turbine shroud assembly including a shroud segment having a leading edge, a trailing edge and a midsection defined therebetween. A shroud support circumferentially spans and supports the shroud segment. The shroud support includes a forward hanger coupled to the leading edge, a midsection hanger coupled to the midsection and an aft hanger coupled to the trailing edge. An annular shroud ring structure includes a midsection position control ring coupled to the midsection hanger and an aft position control ring coupled to the aft hanger. Cooling air is extracted from a compressor positioned upstream of the turbine shroud assembly.Type: GrantFiled: November 30, 2006Date of Patent: June 22, 2010Assignee: General Electric CompanyInventors: Ching-Pang Lee, James Harvey Laflen, Dustin Alfred Placke, George Elliott Moore, Katherine Jaynetorrence Andersen, Daniel Verner Jones
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Patent number: 7740442Abstract: A method for cooling a shroud segment of a gas turbine engine includes providing a turbine shroud assembly including a shroud segment having a leading edge defining a forward face. A turbine nozzle is coupled to the turbine shroud assembly such that a gap is defined between an aft face of an outer band of the turbine nozzle and the forward face, wherein a lip formed on the aft face is positioned radially inwardly with respect to the gap and extends substantially axially downstream from the gap. Cooling air is directed into the gap. Cooling air exiting the gap impinges against the lip to facilitate film cooling the shroud segment.Type: GrantFiled: November 30, 2006Date of Patent: June 22, 2010Assignee: General Electric CompanyInventors: Ching-Pang Lee, Eric Alan Estill, James Harvey Laflen, Daniel Verner Jones
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Patent number: 7722315Abstract: A method of assembling a gas turbine engine includes coupling a turbine shroud assembly within the gas turbine engine. The turbine shroud assembly includes a shroud segment having a leading edge defining a forward face and a radial inner surface. A turbine nozzle is coupled to the turbine shroud assembly such that a gap is defined between an aft face of an outer band of the turbine nozzle and the forward face. A plurality of recuperated cooling openings are defined through the leading edge at an oblique inlet angle with respect to a centerline of the gap and between the forward face and the radial inner surface to direct cooling fluid through the leading edge to facilitate preferential cooling of the leading edge.Type: GrantFiled: November 30, 2006Date of Patent: May 25, 2010Assignee: General Electric CompanyInventors: Ching-Pang Lee, Eric Alan Estill, James Harvey Laflen, Paul Hadley Vitt, Michael Elliot Wymore
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Publication number: 20100111682Abstract: A turbine nozzle includes a row of vanes extending radially between annular outer and inner bands. The outer band includes a pair of radial flanges defining an annular seal groove therebetween. One of the flanges is crenelated to improve nozzle life.Type: ApplicationFiled: October 31, 2008Publication date: May 6, 2010Inventors: Patrick Jarvis Scoggins, James Harvey Laflen, Ching-Pang Lee, Wilson Frost
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Patent number: 7690885Abstract: A method for cooling a shroud segment of a gas turbine engine includes providing a turbine shroud assembly including a shroud segment having a leading edge defining a forward face. A turbine nozzle is coupled to the turbine shroud assembly such that a gap is defined between an aft face of an outer band of the turbine nozzle and the forward face, wherein a lip formed on the aft face is positioned radially inwardly with respect to the gap and extends substantially axially downstream from the gap. Cooling air is directed into the gap. Cooling air exiting the gap impinges against the lip. Post impingement cooling air is directed at the forward face to facilitate forming a film cooling layer on the shroud segment. The film cooling layer is shielded from combustion gases flowing through the gas turbine engine.Type: GrantFiled: November 30, 2006Date of Patent: April 6, 2010Assignee: General Electric CompanyInventors: Ching-Pang Lee, Eric Alan Estill, James Harvey Laflen, Daniel Vern Jones
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Publication number: 20100074745Abstract: A turbine shroud includes a shroud hanger having an arcuate panel from which three inner hooks extend inwardly, and from which two outer hooks extend outwardly therefrom. The two outer hooks effect a statically determinate configuration of the shroud.Type: ApplicationFiled: September 19, 2008Publication date: March 25, 2010Inventors: Daniel Vern Jones, James Harvey Laflen, Richard William Albrecht, Dustin Alfred Placke, Ching-Pang Lee
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Patent number: 7665953Abstract: A method for cooling a shroud segment of a gas turbine engine is provided. The method includes providing a turbine shroud assembly including a shroud segment having an inner surface and a leading edge that is substantially perpendicular to the inner surface, and coupling a turbine nozzle to the turbine shroud segment such that a gap is defined between an aft edge of an outer band of the turbine nozzle and the leading edge. The method also includes directing cooling air into the gap, and directing the cooling air in the gap through at least one cooling hole extending between the leading edge and the inner surface.Type: GrantFiled: November 30, 2006Date of Patent: February 23, 2010Assignee: General Electric CompanyInventors: Ching-Pang Lee, Eric Alan Estill, James Harvey Laflen, Katherine Jaynetorrence Andersen
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Patent number: 7611324Abstract: A method of assembling a gas turbine engine is provided. The method includes coupling at least one turbine nozzle segment within the gas turbine engine. The at least one turbine nozzle segment includes at least one airfoil vane extending between an inner band and an outer band that includes an aft flange and a radial inner surface.Type: GrantFiled: November 30, 2006Date of Patent: November 3, 2009Assignee: General Electric CompanyInventors: Ching-Pang Lee, Eric Alan Estill, James Harvey Laflen, Paul Hadley Vitt, Michael Elliot Wymore
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Patent number: 7604453Abstract: A method for cooling a shroud segment of a gas turbine engine is provided. The method includes providing a turbine shroud assembly including a shroud segment having an inner surface and a leading edge that is substantially perpendicular to the inner surface, and coupling a turbine nozzle to the turbine shroud segment such that a gap is defined between an aft edge of an outer band of the turbine nozzle and the leading edge. The method also includes directing cooling air into the gap, circumferentially mixing the cooling air in a plenum defined within the leading edge to substantially uniformly distribute the cooling air throughout the gap, and directing the cooling air in the gap through at least one cooling hole formed between the plenum and the inner surface.Type: GrantFiled: November 30, 2006Date of Patent: October 20, 2009Assignee: General Electric CompanyInventors: Ching-Pang Lee, Eric Alan Estill, James Harvey Laflen, Katherine Jaynetorrence Andersen
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Publication number: 20080206042Abstract: A method for cooling a shroud segment of a gas turbine engine is provided. The method includes providing a turbine shroud assembly including a shroud segment having an inner surface and a leading edge that is substantially perpendicular to the inner surface, and coupling a turbine nozzle to the turbine shroud segment such that a gap is defined between an aft edge of an outer band of the turbine nozzle and the leading edge. The method also includes directing cooling air into the gap, circumferentially mixing the cooling air in a plenum defined within the leading edge to substantially uniformly distribute the cooling air throughout the gap, and directing the cooling air in the gap through at least one cooling hole formed between the plenum and the inner surface.Type: ApplicationFiled: November 30, 2006Publication date: August 28, 2008Inventors: Ching-Pang Lee, Eric Alan Estill, James Harvey Laflen, Katherine Jaynetorrence Andersen
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Publication number: 20080131260Abstract: A method of assembling a gas turbine engine is provided. The method includes coupling at least one turbine nozzle segment within the gas turbine engine. The at least one turbine nozzle segment includes at least one airfoil vane extending between an inner band and an outer band that includes an aft flange and a radial inner surface. The method also includes coupling at least one turbine shroud segment downstream from the at least one turbine nozzle segment, wherein the at least one turbine shroud segment includes a leading edge and a radial inner surface, and coupling a cooling fluid source in flow communication with the at least one turbine nozzle segment such that cooling fluid channeled to each turbine nozzle outer band aft flange is directed at an oblique discharge angle towards the leading edge of the at least one turbine shroud segment.Type: ApplicationFiled: November 30, 2006Publication date: June 5, 2008Inventors: Ching-Pang Lee, Eric Alan Estill, James Harvey Laflen, Paul Hadley Vitt, Michael Elliot Wymore
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Publication number: 20080131261Abstract: A method of assembling a gas turbine engine is provided. The method includes coupling at least one turbine nozzle segment within the gas turbine engine. The at least one turbine nozzle segment includes at least one airfoil vane extending between an inner band and an outer band that includes an aft flange and a radial inner surface.Type: ApplicationFiled: November 30, 2006Publication date: June 5, 2008Inventors: Ching-Pang Lee, Eric Alan Estill, James Harvey Laflen, Paul Hadley Vitt, Michael Elliot Wymore
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Publication number: 20080131262Abstract: A method for cooling a shroud segment of a gas turbine engine includes providing a turbine shroud assembly including a shroud segment having a leading edge defining a forward face. A turbine nozzle is coupled to the turbine shroud assembly such that a gap is defined between an aft face of an outer band of the turbine nozzle and the forward face, wherein a lip formed on the aft face is positioned radially inwardly with respect to the gap and extends substantially axially downstream from the gap. Cooling air is directed into the gap. Cooling air exiting the gap impinges against the lip to facilitate film cooling the shroud segment.Type: ApplicationFiled: November 30, 2006Publication date: June 5, 2008Inventors: Ching-Pang Lee, Eric Alan Estill, James Harvey Laflen, Daniel Verner Jones
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Publication number: 20080131259Abstract: A method for cooling a shroud segment of a gas turbine engine is provided. The method includes providing a turbine shroud assembly including a shroud segment having an inner surface and a leading edge that is substantially perpendicular to the inner surface, and coupling a turbine nozzle to the turbine shroud segment such that a gap is defined between an aft edge of an outer band of the turbine nozzle and the leading edge. The method also includes directing cooling air into the gap, and directing the cooling air in the gap through at least one cooling hole extending between the leading edge and the inner surface.Type: ApplicationFiled: November 30, 2006Publication date: June 5, 2008Inventors: Ching-Pang Lee, Eric Alan Estill, James Harvey Laflen, Katherine Jaynetorrence Andersen
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Publication number: 20080131263Abstract: A method for cooling a shroud segment of a gas turbine engine includes providing a turbine shroud assembly including a shroud segment having a leading edge defining a forward face. A turbine nozzle is coupled to the turbine shroud assembly such that a gap is defined between an aft face of an outer band of the turbine nozzle and the forward face, wherein a lip formed on the aft face is positioned radially inwardly with respect to the gap and extends substantially axially downstream from the gap. Cooling air is directed into the gap. Cooling air exiting the gap impinges against the lip. Post impingement cooling air is directed at the forward face to facilitate forming a film cooling layer on the shroud segment. The film cooling layer is shielded from combustion gases flowing through the gas turbine engine.Type: ApplicationFiled: November 30, 2006Publication date: June 5, 2008Inventors: Ching-Pang Lee, Eric Alan Estill, James Harvey Laflen, Daniel Verner Jones
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Publication number: 20080127491Abstract: A method of assembling a gas turbine engine includes coupling a turbine shroud assembly within the gas turbine engine. The turbine shroud assembly includes a shroud segment having a leading edge defining a forward face and a radial inner surface. A turbine nozzle is coupled to the turbine shroud assembly such that a gap is defined between an aft face of an outer band of the turbine nozzle and the forward face. A plurality of recuperated cooling openings are defined through the leading edge at an oblique inlet angle with respect to a centerline of the gap and between the forward face and the radial inner surface to direct cooling fluid through the leading edge to facilitate preferential cooling of the leading edge.Type: ApplicationFiled: November 30, 2006Publication date: June 5, 2008Inventors: Ching-Pang Lee, Eric Alan Estill, James Harvey Laflen, Paul Hadley Vitt, Michael Elliot Wymore
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Publication number: 20080131264Abstract: A method for cooling a turbine shroud assembly includes providing a turbine shroud assembly including a shroud segment having a leading edge, a trailing edge and a midsection defined therebetween. A shroud support circumferentially spans and supports the shroud segment. The shroud support includes a forward hanger coupled to the leading edge, a midsection hanger coupled to the midsection and an aft hanger coupled to the trailing edge. An annular shroud ring structure includes a midsection position control ring coupled to the midsection hanger and an aft position control ring coupled to the aft hanger. Cooling air is extracted from a compressor positioned upstream of the turbine shroud assembly.Type: ApplicationFiled: November 30, 2006Publication date: June 5, 2008Inventors: Ching-Pang Lee, James Harvey Laflen, Dustrin Alfred Placke, George Elliott Moore, Katherine Jaynetorrence Andersen, Daniel Verner Jones