Mark Aaron Morris

US Patent:

6481959, Nov 19, 2002

Filed:

Apr 26, 2001

Appl. No.:

09/844889

Inventors:

Mark C. Morris - Phoenix AZ
William J. Howe - Chandler AZ
Carl D. Wright - Chandler AZ
Douglas P. Freiberg - Phoenix AZ
Nnawuihe A. Okpara - Peoria AZ

Assignee:

Honeywell International, Inc. - Morristown NJ

International Classification:

F04D 2938

US Classification:

415115, 415191

Abstract:

A supplemental air cooling system for use in gas turbine engines to inhibit the ingestion of hot flow path gases into circumferential locations of turbine disk cavities is provided. The supplemental air cooling is provided through a simple set of cooling air holes located on each side of the turbine nozzle airfoil trailing edges, and proximately placed to be below the turbine nozzle structural element flow discouragers. Turbine disk cavity cooling purge air entering the disk cavity through the cooling air holes produces dynamic pressure cooling air jets which force the incoming hot ingestion air to turn circumferentially and go back out in the flow path before it enters the turbine disk cavity. The result is a decrease in hot gas ingestion, a reduction in disk rotor and static structural metal temperatures, a reduction in the amount of required cooling air flow, and enhanced performance of the gas turbine engine by virtue of improved specific fuel consumption.

US Patent:

6554562, Apr 29, 2003

Filed:

Jun 15, 2001

Appl. No.:

09/882518

Inventors:

Rodolphe Dudebout - Phoenix AZ
Mark C. Morris - Phoenix AZ
Douglas P. Freiberg - Phoenix AZ
Craig W. McKeever - Gilbert AZ
Richard J. Musiol - Tempe AZ
Ardeshir Riahi - Phoenix AZ
William J. Howe - Chandler AZ

Assignee:

Honeywell International, Inc. - Morristown NJ

International Classification:

F01D 904

US Classification:

415 1, 415191, 4152082

Abstract:

A method and apparatus to reduce the average and maximum temperatures to which the nozzles in the hot-section of gas-turbine engine are subjected is described. The method relates to the circumferential alignment of fuel nozzles and downstream turbine nozzles in a gas turbine engine. This situates the hot-streak emerging from each fuel nozzle in between the like-numbered turbine nozzle airfoils. The most severe operating condition for reducing the durability of nozzle airfoils is the one generating hot operating temperature conditions. By identifying the temperature profile passing through downstream nozzle airfoils, airfoils in static stages can be selectively spaced around the circumference of the ring attached to the casing of the gas turbine engine to avoid high temperature exposure to the airfoils. This method and apparatus mitigates the worst oxidation and thermo-mechanical fatigue damage in the airfoils by allowing the hot gas regions to pass through the path in between two adjacent airfoils.

US Patent:

6931859, Aug 23, 2005

Filed:

Dec 17, 2003

Appl. No.:

10/739754

Inventors:

Peter L. Morgan - Tempe AZ, US
William J. Howe - Chandler AZ, US
Mark C. Morris - Phoenix AZ, US
Adrian R. Allan - Chandler AZ, US
Donald L. Palmer - Cave Creek AZ, US

Assignee:

Honeywell International Inc. - Morristown NJ

International Classification:

F02C007/12

US Classification:

60782, 60785, 137494

Abstract:

Cooling air to the blades and disks of a gas turbine may be modulated to provide a variable turbine cooling flow. A bellows may be extended by providing a high pressure compressor discharge flow to an interior of the bellows. The bellows may be compressed when the interior of the bellows communicates with ambient pressure air. The extension/compression of the bellows moves an arm over orifices in a cooling air flow path. The pressure inside of the bellows is metered to move the arm over at least one orifice, thereby restricting cooling air flow when the engine is running at low power. The pressure inside of the bellows is metered to move the arm to uncover all of the slots to provide maximum cooling flow when the engine is running at high power. The resulting variable cooling flow system results in less need for cooling air at low powers, thus reducing engine fuel consumption.

US Patent:

6942445, Sep 13, 2005

Filed:

Dec 4, 2003

Appl. No.:

10/729695

Inventors:

Mark C. Morris - Phoenix AZ, US
William J. Howe - Chandler AZ, US
George E. Zurmehly - Phoenix AZ, US
Alan G. Tiltman - Fountain Hills AZ, US
George W. Wolfmeyer - Tempe AZ, US
John F. Sciacca - Phoenix AZ, US

Assignee:

Honeywell International Inc. - Morristown NJ

International Classification:

F01D011/08

US Classification:

415 1, 4151731, 415175

Abstract:

A cooled shroud assembly includes an angled slot and a plurality of dilution jet openings. The shroud forward cavity is modified such that at least one recirculation zone is produced. The angled slot forces an axial change in momentum of the hot gas flow and increases radial and axial pressure variation attenuation. The cooled shroud assembly isolates the shroud structure and seals from the hot flow path and a cooling flow from the dilution jet openings dilutes the hot gas flow. A series of recirculation zones shields the shroud carrier and high pressure seals from the hot gas flow.

US Patent:

6974308, Dec 13, 2005

Filed:

Nov 14, 2001

Appl. No.:

09/992250

Inventors:

Steve H. Halfmann - Chandler AZ, US
Yong W. Kim - San Diego CA, US
Mark C. Morris - Phoenix AZ, US
Milton Ortiz - Scottsdale AZ, US
David R. Pack - Plantsville CT, US
Craig A. Wilson - Mesa AZ, US

Assignee:

Honeywell International, Inc. - Morristown NJ

International Classification:

F01D005/16
F01D009/06

US Classification:

416 97R, 415115

Abstract:

A robust multiple-walled, multi-pass, high cooling effectiveness cooled turbine vane or blade designed for ease of manufacturability, minimizes cooling flows on highly loaded turbine rotors. The vane or blade design allows the turbine inlet temperature to increase over current technology levels while simultaneously reducing turbine cooling to low levels. A multi-wall cooling system is described, which meets the inherent conflict to maximize the flow area of the cooling passages while retaining the required section thickness to meet the structural requirements. Independent cooling circuits for the vane or blade's pressure and suction surfaces allow the cooling of the airfoil surfaces to be tailored to specific heat load distributions (that is, the pressure surface circuit is an independent forward flowing serpentine while the suction surface is an independent rearward flowing serpentine). The cooling air for the independent circuits is supplied through separate passages at the base of the vane or blade. The cooling air follows intricate passages to feed the serpentine thin outer wall passages, which incorporate pin fins, turbulators, etc.

US Patent:

6997673, Feb 14, 2006

Filed:

Dec 11, 2003

Appl. No.:

10/734922

Inventors:

Mark C. Morris - Phoenix AZ, US
Steve H. Halfmann - Chandler AZ, US
Craig A. Wilson - Mesa AZ, US
Shawn J. Pollock - Phoenix AZ, US
Carl A. Larson - Sedona AZ, US

Assignee:

Honeywell International, Inc. - Morristown NJ

International Classification:

F01D 25/12
F01D 11/08
F01D 25/24

US Classification:

415 1, 415139, 4151731, 415178, 4152141

Abstract:

A turbine shroud assembly includes forward and aft hangers, an axisymmetric plenum assembly, ceramic shroud segments, ceramic spacers, and forward and aft rope seals. The plenum assembly supplies impingement cooling to the shroud and the hangers. The impingement cooling to the forward and aft hangers is controlled independently to improve blade tip clearance. The rope seals are radially inward from the hangers and reduce cooling flow leakage. The turbine shroud assembly can operate in a higher temperature environment using less cooling flow than the prior art.

US Patent:

7093418, Aug 22, 2006

Filed:

Apr 21, 2004

Appl. No.:

10/830265

Inventors:

Mark C. Morris - Phoenix AZ, US
Alan G. Tiltman - Fountain Hills AZ, US
Kin Poon - Tempe AZ, US
Bradley A. Volkmann - Phoenix AZ, US

Assignee:

Honeywell International, Inc. - Morristown NJ

International Classification:

F02C 7/28
F02C 7/06
F02G 3/00

US Classification:

60 3908, 184 611

Abstract:

A gas turbine engine is configured to use relatively cool, low pressure air discharged from a low pressure compressor to supply buffer air to lubrication sump seals. The engine is further configured such that the lubrication sump is thermally layered by isolating relatively hot, high pressure compressor air from the sump by utilizing a warm vent mixing cavity, which is located radially between of the hot high pressure compressor air and the cool buffer air, which is located in a buffer cavity between the vent cavity and the sump.

US Patent:

7179822, Feb 20, 2007

Filed:

Aug 5, 2003

Appl. No.:

10/634709

Inventors:

Amy Mae Bunker - Ann Arbor MI, US
Mark Anthony Morris - Ypsilanti MI, US
Patrick Michael O'Brien - Stockbridge MI, US
Michael William Wilson - Ann Arbor MI, US

Assignee:

Warner-Lambert Company - Morris Plains NJ
Warner-Lambert Company LLC - Morris Plains NJ

International Classification:

C07D 401/04
C07D 401/14
C07D 403/14
A61K 31/40
A61K 31/44
A61P 19/02

US Classification:

514340, 514341, 514342, 5462681, 5462684

Abstract:

This invention provides compounds defined by Formula I.