Daniel N Demers

US Patent:

6402471, Jun 11, 2002

Filed:

Nov 3, 2000

Appl. No.:

09/706387

Inventors:

Daniel Edward Demers - Ipswich MA
Robert Francis Manning - Newburyport MA
Paul Joseph Acquaviva - Newburyport MA

Assignee:

General Electric Company - Schenectady NY

International Classification:

F01D 518

US Classification:

416 97R, 416 96 R, 416193 A, 415115

Abstract:

A turbine blade includes a platform having an internal cavity formed therein and an airfoil extending radially from the platform. An internal cooling circuit is formed in the airfoil for circulating a coolant therethrough, and at least one supply passage extends between the internal cooling circuit and the internal platform cavity for diverting coolant to the internal platform cavity. The coolant is expelled from holes located in the forward and aft edges of the platform for purging the forward and aft disk wheel spaces and impingement cooling adjacent nozzle bands.

US Patent:

6416284, Jul 9, 2002

Filed:

Nov 3, 2000

Appl. No.:

09/706403

Inventors:

Daniel Edward Demers - Ipswich MA
Robert Francis Manning - Newburyport MA
Paul Joseph Acquaviva - Newburyport MA

Assignee:

General Electric Company - Schenectady NY

International Classification:

F01D 518

US Classification:

416 97R, 416 96 R, 416193 A, 415115

Abstract:

A turbine blade includes a platform and an airfoil extending radially from the platform. An internal cooling circuit is formed in the airfoil for circulating a coolant therethrough to cool the airfoil. At least one supply passage is provided to direct some of the coolant that has passed at least partially through the internal cooling circuit onto the platform for cooling the platform.

US Patent:

6890153, May 10, 2005

Filed:

Apr 29, 2003

Appl. No.:

10/425262

Inventors:

Daniel Edward Demers - Ipswich MA, US
Mohammad Esmaail Taslim - Needham MA, US

Assignee:

General Electric Company - Schenectady NY

International Classification:

F01D005/08

US Classification:

416 97R, 416 96 R, 415 1, 415115

Abstract:

A turbine airfoil includes pressure and suction sidewalls joined together at opposite leading and trailing edges, and at a forward bridge spaced behind the leading edge to define a flow channel. The bridge includes a row of impingement holes. The flow channel includes a row of fins behind the leading edge, a row of first turbulators behind the pressure sidewall, and row of second turbulators behind the suction sidewall. The fins and turbulators have different configurations for increasing internal surface area and heat transfer for back side cooling the leading edge by the cooling air.

US Patent:

6991430, Jan 31, 2006

Filed:

Apr 7, 2003

Appl. No.:

10/408293

Inventors:

Philip Francis Stec - Medford MA, US
Daniel Edward Demers - Ipswich MA, US
Richard Ludwig Schmidt - Marblehead MA, US

Assignee:

General Electric Company - Schenectady NY

International Classification:

F01D 5/18

US Classification:

416 97R, 415115, 4151731, 416 92, 416228

Abstract:

A turbine blade squealer tip has a continuous squealer tip wall extending radially outwardly from and continuously around a tip cap. A recessed tip wall portion of the tip wall is recessed inboard from a pressure side of an airfoil outer wall of an airfoil of the blade forming a tip shelf therebetween. A plurality of film cooling shelf holes are disposed through the tip shelf to an internal cooling circuit of the blade and are spaced away from a junction between the recessed tip wall portion and the tip shelf. The exemplary embodiment of the airfoil includes shelf hole centerlines of the holes passing through pierce points in the shelf. At least a majority of the shelf hole centerlines are angled in outboard directions away from and outboard of the squealer tip wall. A majority of centerlines are angled away from vertical lines passing through the pierce points at first component angles in a range between 2 degrees and 16 degrees.

US Patent:

7377742, May 27, 2008

Filed:

Oct 14, 2005

Appl. No.:

11/250660

Inventors:

Jason David Shapiro - Methuen MA, US
Daniel Demers - Ipswich MA, US
Robert Patrick Tameo - Peabody MA, US
Tyler F. Hooper - Amesbury MA, US
Robert Alexander Nicoll - Beverly MA, US
Douglas Patrick Probasco - Peabody MA, US

Assignee:

General Electric Company - Schenectady NY

International Classification:

F01D 25/26

US Classification:

415108, 415115, 4151821

Abstract:

A method for assembling a gas turbine engine includes coupling a rotor assembly including a plurality of rotor blades about a rotatable main shaft of the gas turbine engine. The main shaft is aligned in an axial direction of the gas turbine engine. A shroud assembly is coupled to the gas turbine engine. The shroud assembly includes a plurality of shroud segments circumferentially coupled about the rotor assembly such that a shroud spacing gap is formed in the axial direction between adjacent shroud segments. A cooling fluid source is coupled to each shroud segment such that cooling fluid is channeled through each shroud segment into a corresponding shroud spacing gap to facilitate positive purge flow through the shroud spacing gap.

US Patent:

20090293495, Dec 3, 2009

Filed:

May 29, 2008

Appl. No.:

12/129375

Inventors:

Victor Hugo Silva Correia - Milton Hills NH, US
Daniel Edward Demers - Ipswich MA, US
Robert Francis Manning - Newburyport MA, US

Assignee:

General Electric Company - Schenectady NY

International Classification:

F02C 7/18
F01D 9/02

US Classification:

60782, 415115

Abstract:

A turbine airfoil for a gas turbine engine includes: (a) spaced-apart pressure and suction sidewalls extending between a leading edge and a trailing edge; (b) a first cavity disposed between the pressure and suction sidewalls, the first cavity being adapted to be fed cooling air from a source within the engine, and connected to at least one film cooling hole which communicates with an exterior surface of the airfoil; (c) a second cavity disposed between the pressure and suction sidewalls, the second cavity being adapted to be fed cooling air from a source within the engine, and connected to at least one film cooling hole which communicates solely with the suction sidewall of the airfoil; and (d) a metering structure adapted to substantially restrict air flow into the second cavity.

US Patent:

55557210, Sep 17, 1996

Filed:

Sep 28, 1994

Appl. No.:

8/313951

Inventors:

John J. Bourneuf - Jamaica Plain MA
Dean T. Lenahan - Cinncinnati OH
Daniel E. Demers - Ipswich MA
Larry W. Plemmons - Fairfield OH

Assignee:

General Electric Company - Cincinnati OH

International Classification:

F02C 712

US Classification:

60 3975

Abstract:

A turbine cooling supply circuit for a gas turbine engine is disclosed in which the flow of coolant through the engine is directed to minimize temperature rise prior to discharge into the turbine. Bleed air from an impeller stage of a compressor is channeled along a backwall thereof into a cavity disposed radially inwardly of a combustor casing. The cavity is divided by a lightweight tubular member having a circumferentially uniform contour into a non-flow zone proximate the casing and a flow zone proximate a rotor. Maintenance of a high tangential flow component in the bleed flow in the flow zone facilitates discharge of the flow through apertures in a rotor into a bore portion thereof, minimizing work input and temperature rise of the coolant. Further, the coolant is isolated from the hot casing wall and nonuniform structures disposed thereon, thereby reducing heat flux and temperature rise associated with aerodynamic drag caused by increased surface area, nonuniform interruptions in the stationary surface or any combination thereof. Features are incorporated in the tubular member to provide margin against buckling.

US Patent:

61831980, Feb 6, 2001

Filed:

Nov 16, 1998

Appl. No.:

9/192229

Inventors:

Robert F. Manning - Newburyport MA
Paul J. Acquaviva - Wakefield MA
Daniel E. Demers - Ipswich MA

Assignee:

General Electric Company - Cincinnati OH

International Classification:

F01D 518

US Classification:

416 97R

Abstract:

A gas turbine engine airfoil includes first and second sidewalls joined together at opposite leading and trailing edges, and spaced apart from each other therebetween to define a leading edge channel extending longitudinally from a root to a tip of the airfoil. A plurality of film cooling holes extend through the leading edge and are disposed in flow communication with the leading edge channel. An isolation plenum extends along the first sidewall and adjacent the leading edge channel, and is separated therefrom by a partition having a plurality of inlet holes. A plurality of film cooling gill holes extend through the first sidewall, and are disposed in flow communication with the isolation plenum. Cooling air is channeled from the leading edge channel to the isolation plenum for feeding the gill holes with reduced pressure air.