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Ardeshir Riahi

from Phoenix, AZ
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Also known as:
Ardeshir Lasalle, A Riahi, Riahi Ardeshire
Phone and address:
4717 Beverly Rd, Phoenix, AZ 85032
Connected to:
Behnam R Riahi, 38Giti Riahi, 82Omar M Riahi, 42Sara Riahi, 28Siavash Riahi, 84Marc A Riak, 60Allison L Rial, 35Andres J Rial, 48Antonio J Rial, 55Christina Rial, 39

Ardeshir Riahi Phones & Addresses

  • 4717 Beverly Rd, Phoenix, AZ 85032
  • 3104 Amber Ridge Way, Phoenix, AZ 85048
  • 13625 48Th St, Phoenix, AZ 85044
  • Paradise Vly, AZ
  • 6803 E North Ln, Paradise Valley, AZ 85253
  • Scottsdale, AZ

Work

Company: Honeywell Apr 1996 to Jun 2006 Position: Principal engineer

Education

Degree: Doctorates, Doctor of Philosophy School / High School: The University of British Columbia 1983 to 1990 Specialities: Philosophy, Mechanical Engineering

Skills

Aerospace • Engineering • Engineering Management • Mechanical Engineering • Aerospace Engineering • Finite Element Analysis • Six Sigma • Systems Engineering • Ansys • Matlab • Continuous Improvement • Manufacturing

Languages

English • Farsi

Industries

Aviation & Aerospace

Resumes

Resumes

Ardeshir Riahi Photo 1

Senior Technical Manager

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Location:
Phoenix, AZ
Industry:
Aviation & Aerospace
Work:
Honeywell Apr 1996 - Jun 2006
Principal Engineer
Honeywell Apr 1996 - Jun 2006
Senior Technical Manager
Pratt & Whitney Canada Apr 1990 - Apr 1996
Principal Engineer
Bc Research Centre Feb 1988 - Nov 1989
Consultant Engineer
Education:
The University of British Columbia 1983 - 1990
Doctorates, Doctor of Philosophy, Philosophy, Mechanical Engineering The University of British Columbia 1983 - 1985
Masters, Mechanical Engineering Chatrapati Sahuji Maharaj Kanpur University, Kanpur 1980 - 1983
Bachelors, Bachelor of Science, Mechanical Engineering
Skills:
Aerospace
Engineering
Engineering Management
Mechanical Engineering
Aerospace Engineering
Finite Element Analysis
Six Sigma
Systems Engineering
Ansys
Matlab
Continuous Improvement
Manufacturing
Languages:
English
Farsi

Publications

Us Patents

Gas Turbine Combustor Heat Shield Impingement Cooling Baffle

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US Patent:
6792757, Sep 21, 2004
Filed:
Nov 5, 2002
Appl. No.:
10/288950
Inventors:
Frederick G. Borns - Chandler AZ
Ardeshir Riahi - Phoenix AZ
Assignee:
Honeywell International Inc. - Morristown NJ
International Classification:
F02G 100
US Classification:
60772, 60 3911, 60752
Abstract:
A heat shield for a combustor dome includes U-shaped baffles on the outer diameter area of the upstream surface of the heat shield. The baffles are clocked with respect to the impingement openings in the combustor dome. The baffles increase cooling of the heat shield by segregating the cooling air flow from the impingement openings and by reducing cross-flow at the outer diameter of the heat shield. The baffles also function as heat shield stiffeners. Slots extend radially inward from the outer rim of the heat shield. Keyholes are at the inner ends of the slots. The slots and keyholes reduce the hoop stresses of the heat shield.

Gas Turbine Engine Including Airfoils Having An Improved Airfoil Film Cooling Configuration And Method Therefor

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US Patent:
7223072, May 29, 2007
Filed:
Jan 27, 2004
Appl. No.:
10/766231
Inventors:
Ardeshir Riahi - Phoenix AZ, US
Robert McDonald - Chandler AZ, US
Frederick G. Borns - Chandler AZ, US
Assignee:
Honeywell International, Inc. - Morristown NJ
International Classification:
F01D 5/18
US Classification:
416 61, 416 97 R, 29557, 29889721
Abstract:
An airfoil for a gas turbine engine blade includes a plurality of film cooling holes extending through its outer surface. The film cooling holes are formed by defining at least a first datum structure and a second datum structure, and then forming each film cooling hole at a location on the airfoil outer surface relative to the first and second datum structures. As a result, each film cooling hole has a centerline extending therethrough that forms a compound angle with respect to a tangent to the outer surface, and the distance between the centerlines of each film cooling hole is at least a predetermined minimum distance.

Turbine Rotor Cooling Flow System

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US Patent:
8277169, Oct 2, 2012
Filed:
Jun 16, 2005
Appl. No.:
11/155399
Inventors:
Ardeshir Riahi - Scottsdale AZ, US
Frederick G. Borns - Chandler AZ, US
Vivek Agarwal - Chandler AZ, US
Assignee:
Honeywell International Inc. - Morristown NJ
International Classification:
F01D 5/08
US Classification:
415115, 415151, 415176, 416 96 R, 416 97 R
Abstract:
Disclosed herein is an apparatus comprising a disk coverplate for a turbine rotor, the disk coverplate comprising a plurality of cooling holes, wherein the distance between the centers of any two adjacent cooling holes is greater than twice the average diameter of the two adjacent cooling holes. A method to control turbine cooling air flow is also disclosed.

Air Cooled Turbine Blades And Methods Of Manufacturing

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US Patent:
8292581, Oct 23, 2012
Filed:
Jan 9, 2008
Appl. No.:
11/971459
Inventors:
Kin C. Poon - Tempe AZ, US
Malak F. Malak - Tempe AZ, US
Rajiv Rana - Tempe AZ, US
Ardeshir Riahi - Scottsdale AZ, US
David H. Chou - Phoenix AZ, US
Assignee:
Honeywell International Inc. - Morristown NJ
International Classification:
F01D 5/08
US Classification:
416 97R
Abstract:
An air-cooled turbine blade and methods of manufacturing the blade are provided. The blade includes a suction side flow circuit formed within its interior and defined at least by an interior surface of a convex suction side wall, a pressure side flow circuit formed within the blade interior and defined at least by an interior surface of a concave pressure side wall, and a center flow circuit including a first section and a second section, the first section disposed between the suction side flow circuit and the pressure side flow circuit, and the second section in flow communication with the first section and a plurality of openings of a leading edge wall and defined at least partially by an interior surface of the leading edge wall.

Turbine Blade Assemblies And Methods Of Manufacturing The Same

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US Patent:
8292587, Oct 23, 2012
Filed:
Dec 18, 2008
Appl. No.:
12/338746
Inventors:
Kin Poon - Tempe AZ, US
Rajiv Rana - Phoenix AZ, US
Bob Mitlin - Scottsdale AZ, US
Ardeshir Riahi - Scottsdale AZ, US
David Chou - Phoenix AZ, US
Steve Halfmann - Chandler AZ, US
Frank Mignano - Phoenix AZ, US
Assignee:
Honeywell International Inc. - Morristown NJ
International Classification:
F01D 11/00
US Classification:
416193A, 415115, 415116, 416 90 R, 416 95, 416 96 R, 416 97 R, 416193 R
Abstract:
A turbine blade assembly includes an airfoil, a platform, and a first cover plate. A center flow path extends through the platform in communication with an internal cooling circuit of the airfoil, which extends from a first side of the platform. A second side of the platform is located opposite the platform from the first side. An edge of the platform extends between the first and second sides and, a first passage is formed between the first and second sides and includes a first inlet and a first outlet. The first passage extends from the center flow path toward the platform edge, and a first groove is formed on the second side of the platform and extends from the first outlet of the first passage toward the edge of the platform. The first cover plate is disposed over the second side of the platform covering the first groove.

Turbine Blades And Methods Of Forming Modified Turbine Blades And Turbine Rotors

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US Patent:
8297935, Oct 30, 2012
Filed:
Nov 18, 2008
Appl. No.:
12/273108
Inventors:
Bob Mitlin - Scottsdale AZ, US
Mark C. Morris - Phoenix AZ, US
Steve Halfmann - Chandler AZ, US
Ardeshir Riahi - Scottsdale AZ, US
Assignee:
Honeywell International Inc. - Morristown NJ
International Classification:
F01D 5/14
US Classification:
416243, 416226, 416223 R, 416 97 R, 416219 R, 298897, 2940701
Abstract:
Turbine blades and methods of forming modified turbine blades and turbine rotors for use in an engine are provided. In an embodiment, by way of example only, a turbine blade includes a platform and an airfoil. The platform includes a surface configured to define a portion of a flowpath, and the surface includes an initial contour configured to plastically deform into an intended final contour after an initial exposure of the blade to an operation of the engine. The airfoil extends from the platform.

Turbine Blades And Methods Of Manufacturing

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US Patent:
8206108, Jun 26, 2012
Filed:
Dec 10, 2007
Appl. No.:
11/953489
Inventors:
Ardeshir Riahi - Scottsdale AZ, US
Kin Poon - Tempe AZ, US
David Chou - Phoenix AZ, US
Malak F. Malak - Tempe AZ, US
Assignee:
Honeywell International Inc. - Morristown NJ
International Classification:
F01D 5/18
US Classification:
416 97R, 416228
Abstract:
A turbine blade includes a convex suction side wall, a concave pressure side wall, a tip wall, an internal cooling circuit, and a plurality of tip edge channels. The tip wall is recessed from a first tip edge of the suction side wall and a second tip edge of the concave pressure side wall to define a suction side wall tip section and a pressure side wall tip section, and the suction side wall tip section is shorter than the pressure side wall tip section. The internal cooling circuit is formed at least partially between the convex suction side wall, the concave pressure side wall, and the tip wall. The plurality of tip edge channels formed through the first tip edge of the convex suction side wall extend to the internal cooling circuit. Methods of manufacturing turbine blades are also provided.

Combustor Hot Streak Alignment For Gas Turbine Engine

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US Patent:
20030002975, Jan 2, 2003
Filed:
Jun 15, 2001
Appl. No.:
09/882518
Inventors:
Rodolphe Dudebout - Phoenix AZ, US
Mark Morris - Phoenix AZ, US
Douglas Freiberg - Phoenix AZ, US
Craig McKeever - Gilbert AZ, US
Richard Musiol - Tempe AZ, US
Ardeshir Riahi - Phoenix AZ, US
William Howe - Chandler AZ, US
Assignee:
Honeywell International, Inc. - Morristown NJ
International Classification:
F03D001/00
US Classification:
415/001000
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.
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Ardeshir Riahi from Phoenix, AZ Get Report