Jeff Howe

US Patent:

8147197, Apr 3, 2012

Filed:

Mar 10, 2009

Appl. No.:

12/401037

Inventors:

Mark C. Morris - Phoenix AZ, US
Jason Smoke - Phoenix AZ, US
Jeff Howe - Chandler AZ, US
Bob Mitlin - Scottsdale AZ, US

Assignee:

Honeywell International, Inc. - Morristown NJ

International Classification:

F01D 5/18

US Classification:

416 97R, 416223 R, 416233, 416500

Abstract:

A turbine blade assembly is provided. The turbine blade assembly comprises a turbine blade comprising a cavity, and a blade platform supporting the turbine blade, the cavity extending into the blade platform. The blade platform comprises an upper surface adjacent the turbine blade and a lower surface comprising a first rib, the cavity extending into the first rib, the first rib coupled to the lower surface, tapering as it extends away from the turbine blade, and comprising a first port extending from the cavity to the upper surface.

US Patent:

8167557, May 1, 2012

Filed:

Aug 7, 2008

Appl. No.:

12/187791

Inventors:

Tom Elliott - Chandler AZ, US
Greg Heitland - Mesa AZ, US
Jeff Howe - Chandler AZ, US
Mark C. Morris - Phoenix AZ, US

Assignee:

Honeywell International Inc. - Morristown NJ

International Classification:

F01D 5/18

US Classification:

416 97R, 416 92, 416 96 R, 416235, 415 17, 415115, 415119, 415191

Abstract:

A gas turbine engine assembly has combustion gases flowing through a gas flow path. The gas turbine engine assembly includes a stator assembly comprising a stator vane that extends into the gas flow path; and a turbine rotor assembly downstream of the stator assembly and comprising a turbine platform and a turbine rotor blade extending from the turbine platform into the mainstream combustion gases flow path. The turbine rotor blade includes a pressure side and a suction side opposing the pressure side that extend from a leading edge to a trailing edge. The combustion gases form horseshoe vortices at a formation area adjacent the leading edge of the turbine rotor blade, and the turbine rotor assembly further includes a first set of holes in the turbine platform for directing first jets into the formation area of the horseshoe vortices.

US Patent:

8262342, Sep 11, 2012

Filed:

Jul 10, 2008

Appl. No.:

12/171000

Inventors:

Mark C. Morris - Phoenix AZ, US
Jeff Howe - Chandler AZ, US
Greg Heitland - Mesa AZ, US
Jason Smoke - Phoenix AZ, US

Assignee:

Honeywell International Inc. - Morristown NJ

International Classification:

F01D 5/14

US Classification:

415115, 415191

Abstract:

A gas turbine engine assembly includes a housing including an annular duct wall that at least partially defines a mainstream hot gas flow path configured to receive mainstream hot gas flow. The assembly further includes a stator assembly including a stator vane that extends into the mainstream hot gas flow path and a turbine rotor assembly downstream of the stator assembly that includes a turbine disk and a turbine blade extending from the turbine disk into the mainstream hot gas flow path. The stator assembly and turbine assembly define a turbine disk cavity, and the turbine disk cavity includes a recirculation cavity configured to recirculate gas ingested from the mainstream hot gas flow path back into the mainstream hot gas flow path.

US Patent:

8206080, Jun 26, 2012

Filed:

Jun 12, 2008

Appl. No.:

12/138108

Inventors:

Jeff Howe - Chandler AZ, US
Mark C. Morris - Phoenix AZ, US
Khosro Molla Hosseini - Phoenix AZ, US

Assignee:

Honeywell International Inc. - Morristown NJ

International Classification:

F01D 25/00

US Classification:

415 584, 415177, 4151995

Abstract:

A gas turbine engine includes a housing with a duct wall that defines a generally annular and axially elongated hot gas flow path for passage of combustion gas. The engine further includes a contoured shroud mounted within the internal engine cavity and defining a hot gas recirculation pocket for receiving hot gas ingested from the hot gas flow path through the annular space and for recirculating the ingested hot gas back through the annular space to the hot gas flow path. The contoured shroud includes a base wall extending radially inwardly from the duct wall, an inboard wall extending from the base wall in an axial direction toward the rotor, and an end wall extending from the inboard wall in a radially outward direction. The end wall terminates in a circumferentially extending free edge disposed in proximity to the annular space. The end wall defines an opening.

US Patent:

20110299972, Dec 8, 2011

Filed:

Jun 4, 2010

Appl. No.:

12/794433

Inventors:

Mark C. Morris - Phoenix AZ, US
Alexander MirzaMoghadam - Phoenix AZ, US
Khosro Molla Hosseini - Scottsdale AZ, US
Kin Poon - Tempe AZ, US
Jeff Howe - Chandler AZ, US
Alan G. Tiltman - Fountain Hills AZ, US

Assignee:

HONEYWELL INTERNATIONAL INC. - Morristown NJ

International Classification:

F01D 1/00

US Classification:

415 1, 4151821

Abstract:

An impeller or axial stage compressor disk backface shroud for use with a gas turbine engine is disclosed. The backface shroud includes, but is not limited to, a substantially funnel shaped body having a surface. The substantially funnel shaped body is configured to be statically mounted to the gas turbine engine substantially coaxially with the impeller or axial stage compressor disk. The surface and a backface of the impeller or axial stage compressor disk form a cavity that guides an airflow portion to a turbine when the substantially funnel shaped body is mounted coaxially with the impeller or axial stage compressor disk and axially spaced apart therefrom. The airflow portion has a tangential velocity and a recessed groove in the surface of the backface shroud is oriented generally transversely to the tangential velocity to at least partially interfere with the airflow portion, thus affecting static pressure in the cavity.

US Patent:

20130051974, Feb 28, 2013

Filed:

Aug 25, 2011

Appl. No.:

13/218224

Inventors:

Kin Poon - Tempe AZ, US
Jeff Howe - Chandler AZ, US
Michael T. Todd Barton - Phoenix AZ, US
Frank Mignano - Phoenix AZ, US
Mahmoud Mansour - Phoenix AZ, US
Nick Nolcheff - Chandler AZ, US
Scott Taylor - Phoenix AZ, US
Steve Trzcinski - Gilbert AZ, US
Ardeshir Riahi - Scottsdale AZ, US
Jeff Guymon - Gilbert AZ, US
Alan Hemmingson - Tempe AZ, US
Alex Mirzamoghadam - Phoenix AZ, US

Assignee:

HONEYWELL INTERNATIONAL INC. - Morristown NJ

International Classification:

F04D 27/02
F01D 5/08

US Classification:

415 1, 415180

Abstract:

Gas turbine engines and methods for cooling components thereof with mid-impeller bleed (MIB) cooling air having a pressure are provided. The gas turbine engine has a compressor comprising an impeller body and an impeller shroud at least partially surrounding the impeller body. The impeller shroud has a plurality of MIB openings disposed therein. At least one edge treatment is provided thereto. The edge treatment substantially preserves pressure of the cooling air during entrance into and discharge out of the MIB opening. The plurality of MIB openings may be extended MIB openings in a thickened impeller shroud. The centerline of the MIB openings may be oriented to be substantially aligned with an averaged local absolute flow velocity vector of the cooling air at the inlet section of the MIB opening in order to extract cooling air in a direction that has a vector component in a tangential, an axial, and a radial flow direction.

US Patent:

20180274377, Sep 27, 2018

Filed:

Mar 27, 2017

Appl. No.:

15/470471

Inventors:

- Morris Plains NJ, US
Daniel C. Crites - Mesa AZ, US
Jeff Howe - Chandler AZ, US

Assignee:

HONEYWELL INTERNATIONAL INC. - Morris Plains NJ

International Classification:

F01D 9/04
F01D 9/06

Abstract:

Vane impingement tubes having blockage deterrent features are provided, as turbine nozzles containing blockage-resistant vane impingement tubes. In an embodiment, the turbine nozzle includes inner and outer annular endwalls, and turbine nozzle vanes arranged in an annular array between the outer and inner annular endwalls. Vane impingement tubes are inserted into the turbine nozzle vanes. The vane impingement tubes each includes a tube body, an impingement outlet formed in the tube body and configured to discharge airflow for impingement against one of the turbine nozzle vanes, a first flow-turning feature located in the tube body, and an inlet formed in the tube body and configured receive cooling airflow in a substantially radial direction. The first flow-turning feature is shaped and positioned to turn the airflow received through the inlet in a substantially axial direction, which is perpendicular to the radial direction, prior to discharge through the impingement outlet.

US Patent:

20150354461, Dec 10, 2015

Filed:

Apr 3, 2014

Appl. No.:

14/244219

Inventors:

- Morristown NJ, US
Daniel C. Crites - Mesa AZ, US
Yogendra Yogi Sheoran - Scottsdale AZ, US
Mark C. Morris - Phoenix AZ, US
Jeff Howe - Chandler AZ, US
Ardeshir Riahi - Scottsdale AZ, US

Assignee:

HONEYWELL INTERNATIONAL INC. - Morristown NJ

International Classification:

F02C 7/30
F02C 7/052
F02C 3/04
F02C 7/18

Abstract:

A cooling arrangement is provided for a gas turbine engine with a turbine section. The cooling arrangement includes a first conduit to receive cooling air that includes particles; a separator system coupled to the first conduit to receive the cooling air and configured to remove at least a portion of the particles to result in relatively clean cooling air and scavenge air including the portion of the particles; and a second conduit coupled to the separator system and configured to direct the relatively clean cooling air to the turbine section.