Patents by Inventor Mohammed Ennacer
Mohammed Ennacer 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: 11578600Abstract: A turbine blade for a gas turbine engine has an airfoil including leading and trailing edges joined by spaced-apart pressure and suction sides to provide an exterior airfoil surface extending from a platform in a spanwise direction to a tip. The external airfoil surface is formed in substantial conformance with multiple cross-sectional profiles of the airfoil described by a set of Cartesian coordinates set forth in Table 1, the Cartesian coordinates provided by an axial coordinate scaled by a local axial chord, a circumferential coordinate scaled by a local axial chord, and a span location.Type: GrantFiled: October 15, 2021Date of Patent: February 14, 2023Assignee: PRATT & WHITNEY CANADA CORP.Inventors: Marc Tardif, Francois Roy, Jonathan Dorai, Mohammed Ennacer, Benjamin Luymes
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Patent number: 11536141Abstract: A turbine vane for a gas turbine engine has an airfoil including leading and trailing edges joined by spaced-apart pressure and suction sides to provide an external airfoil surface. The surface is formed in substantial conformance with multiple cross-sectional profiles of the airfoil defined by a set of Cartesian coordinates set forth in Table 1, the Cartesian coordinates provided by an axial coordinate scaled by a local axial chord, a circumferential coordinate scaled by a local axial chord, and a span location.Type: GrantFiled: February 4, 2022Date of Patent: December 27, 2022Assignee: PRATT & WHITNEY CANADA CORP.Inventors: Mohammed Ennacer, Panagiota Tsifourdaris, Remo Marini, Niloofar Moradi, Daniel Lecuyer, Patricia Phutthavong, Farzad Ashrafi
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Patent number: 11377965Abstract: A component for a gas turbine engine includes, among other things, an airfoil that extends between a leading edge and a trailing edge and a cooling circuit disposed inside of the airfoil. The cooling circuit includes at least one core cavity that extends inside of the airfoil, a baffle received within the at least one core cavity, a plurality of pedestals positioned adjacent to the at least one core cavity and a first plurality of axial ribs positioned between the plurality of pedestals and the trailing edge of the airfoil.Type: GrantFiled: March 6, 2017Date of Patent: July 5, 2022Assignee: Raytheon Technologies CorporationInventors: Shawn J. Gregg, Dominic J. Mongillo, Michael Leslie Clyde Papple, Russell J. Bergman, Mohammed Ennacer
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Patent number: 11365645Abstract: A turbine shroud segment for a gas turbine engine having an annular gas path extending about an engine axis. The engine has a turbine rotor mounted for rotation about the axis and having a plurality of blades extending into the gas path. The turbine shroud segment includes a body extending axially between a leading edge and a trailing edge and circumferentially between a first and second lateral edges. The body has a radially outer surface and a radially inner surface. The radially outer surface includes a textured surface exposed to a cooling flow. The radially inner surface defines an outer flow boundary surface of the gas path next to a tip of one of the blades. A cooling flow passageway is defined in the body and extends axially between one or more cooling inlets receiving the cooling flow from the textured surface and one or more cooling outlets.Type: GrantFiled: October 7, 2020Date of Patent: June 21, 2022Assignee: PRATT & WHITNEY CANADA CORP.Inventors: Jacob Biernat, Mohammed Ennacer, Remy Synnott, Andrey Potiforov
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Publication number: 20220106887Abstract: A turbine shroud segment for a gas turbine engine having an annular gas path extending about an engine axis. The engine has a turbine rotor mounted for rotation about the axis and having a plurality of blades extending into the gas path. The turbine shroud segment includes a body extending axially between a leading edge and a trailing edge and circumferentially between a first and second lateral edges. The body has a radially outer surface and a radially inner surface. The radially outer surface includes a textured surface exposed to a cooling flow. The radially inner surface defines an outer flow boundary surface of the gas path next to a tip of one of the blades. A cooling flow passageway is defined in the body and extends axially between one or more cooling inlets receiving the cooling flow from the textured surface and one or more cooling outlets.Type: ApplicationFiled: October 7, 2020Publication date: April 7, 2022Inventors: Jacob BIERNAT, Mohammed ENNACER, Remy SYNNOTT, Andrey POTIFOROV
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Patent number: 11274569Abstract: A turbine shroud segment has a body extending axially between a leading edge and a trailing edge and circumferentially between a first and a second lateral edge. Upstream and downstream plenums are defined in the body. The upstream plenum has a plurality of cooling inlets. The downstream plenum has a plurality of cooling outlets. A flow constricting slot extends across the body between the first and second lateral edges. The flow constricting slot fluidly connects the downstream plenum to the upstream plenum.Type: GrantFiled: October 24, 2019Date of Patent: March 15, 2022Assignee: PRATT & WHITNEY CANADA CORP.Inventors: Mohammed Ennacer, Ion Dinu
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Patent number: 11118475Abstract: A turbine shroud segment has a body having a radially outer surface and a radially inner surface extending axially between a leading edge and a trailing edge and circumferentially between a first and a second lateral edge. A first serpentine channel is disposed axially along the first lateral edge. A second serpentine channel is disposed axially along the second lateral edge. The first and second serpentine channels each define a tortuous path including axially extending passages between a front inlet proximate the leading edge and a rear outlet at the trailing edge of the body.Type: GrantFiled: January 23, 2020Date of Patent: September 14, 2021Assignee: PRATT & WHITNEY CANADA CORP.Inventors: Remy Synnott, Mohammed Ennacer, Chris Pater, Denis Blouin, Kapila Jain, Farough Mohammadi
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Patent number: 11015450Abstract: A two-stage high pressure turbine includes a second stage blade having an airfoil with a profile substantially in accordance with at least an intermediate portion of the Cartesian coordinate values of X, Y and Z set forth in Table 2. The X and Y values are distances, which when smoothly connected by an appropriate continuing curve, define airfoil profile sections at each distance Z. The profile sections at each distance Z are joined smoothly to one another to form a complete airfoil shape.Type: GrantFiled: June 14, 2019Date of Patent: May 25, 2021Assignee: PRATT & WHITNEY CANADA CORP.Inventors: Mohammed Ennacer, Dan Olaru, Jasrobin Grewal, Gaetan Girard
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Patent number: 10941709Abstract: The gas turbine engine can have a gas path extending in serial flow communication across a compressor, a combustion chamber, and a turbine, the turbine having at least one multistage turbine section having a front toward the combustion chamber and a rear opposite the front, a plenum radially outward of the gas path, and a plurality of aperture sets interspaced from one another between the front and the rear, the aperture sets providing fluid flow communication from the plenum to the gas path, and a cooling air path having an outlet fluidly connected to the plenum at the rear of the plenum.Type: GrantFiled: September 28, 2018Date of Patent: March 9, 2021Assignee: PRATT & WHITNEY CANADA CORP.Inventors: Ion Dinu, Mohammed Ennacer, René Paquet
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Publication number: 20200392850Abstract: A two-stage high pressure turbine includes a second stage blade having an airfoil with a profile substantially in accordance with at least an intermediate portion of the Cartesian coordinate values of X, Y and Z set forth in Table 2. The X and Y values are distances, which when smoothly connected by an appropriate continuing curve, define airfoil profile sections at each distance Z. The profile sections at each distance Z are joined smoothly to one another to form a complete airfoil shape.Type: ApplicationFiled: June 14, 2019Publication date: December 17, 2020Inventors: Mohammed ENNACER, Dan OLARU, Jasrobin GREWAL, Gaetan GIRARD
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Publication number: 20200277876Abstract: A turbine shroud segment has a body having a radially outer surface and a radially inner surface extending axially between a leading edge and a trailing edge and circumferentially between a first and a second lateral edge. A first serpentine channel is disposed axially along the first lateral edge. A second serpentine channel is disposed axially along the second lateral edge. The first and second serpentine channels each define a tortuous path including axially extending passages between a front inlet proximate the leading edge and a rear outlet at the trailing edge of the body.Type: ApplicationFiled: January 23, 2020Publication date: September 3, 2020Inventors: Remy SYNNOTT, Mohammed ENNACER, Chris PATER, Denis BLOUIN, Kapila JAIN, Farough MOHAMMADI
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Publication number: 20200149429Abstract: A turbine shroud segment has a body extending axially between a leading edge and a trailing edge and circumferentially between a first and a second lateral edge. Upstream and downstream plenums are defined in the body. The upstream plenum has a plurality of cooling inlets. The downstream plenum has a plurality of cooling outlets. A flow constricting slot extends across the body between the first and second lateral edges. The flow constricting slot fluidly connects the downstream plenum to the upstream plenum.Type: ApplicationFiled: October 24, 2019Publication date: May 14, 2020Inventors: Mohammed ENNACER, Ion DINU
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Publication number: 20200102887Abstract: The gas turbine engine can have a gas path extending in serial flow communication across a compressor, a combustion chamber, and a turbine, the turbine having at least one multistage turbine section having a front toward the combustion chamber and a rear opposite the front, a plenum radially outward of the gas path, and a plurality of aperture sets interspaced from one another between the front and the rear, the aperture sets providing fluid flow communication from the plenum to the gas path, and a cooling air path having an outlet fluidly connected to the plenum at the rear of the plenum.Type: ApplicationFiled: September 28, 2018Publication date: April 2, 2020Inventors: Ion DINU, Mohammed ENNACER, René PAQUET
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Patent number: 10570773Abstract: A turbine shroud segment has a body having a radially outer surface and a radially inner surface extending axially between a leading edge and a trailing edge and circumferentially between a first and a second lateral edge. A first serpentine channel is disposed axially along the first lateral edge. A second serpentine channel is disposed axially along the second lateral edge. The first and second serpentine channels each define a tortuous path including axially extending passages between a front inlet proximate the leading edge and a rear outlet at the trailing edge of the body.Type: GrantFiled: December 13, 2017Date of Patent: February 25, 2020Assignee: Pratt & Whitney Canada Corp.Inventors: Remy Synnott, Mohammed Ennacer, Chris Pater, Denis Blouin, Kapila Jain, Farough Mohammadi
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Patent number: 10533454Abstract: A turbine shroud segment comprises a body having an upstream end portion and a downstream end portion relative to a flow of gases through the gas path. A core cavity is defined in the body and extends axially from the upstream end portion to the downstream end portion. A plurality of cooling inlets is defined in the upstream end portion of the body for feeding coolant in the core cavity. A plurality of cooling outlets is defined in the downstream end portion of the body for discharging coolant from the core cavity. Pedestals are provided in the core cavity.Type: GrantFiled: December 13, 2017Date of Patent: January 14, 2020Assignee: Pratt & Whitney Canada Corp.Inventors: Remy Synnott, Mohammed Ennacer, Chris Pater, Denis Blouin, Kapila Jain, Farough Mohammadi
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Patent number: 10502093Abstract: A turbine shroud segment has a body extending axially between a leading edge and a trailing edge and circumferentially between a first and a second lateral edge. A core cavity is defined in the body and extends axially from a front end adjacent the leading edge to a rear end adjacent to the trailing edge. A plurality of cooling inlets and outlets are respectively provided along the front end and the rear end of the core cavity. A crossover wall extends across the core cavity and defines a row of crossover holes configured to accelerate the flow of coolant directed into the core cavity via the cooling inlets. The crossover wall is positioned to accelerate the coolant flow at the beginning of the cooling scheme where the shroud segment is the most thermally solicited.Type: GrantFiled: December 13, 2017Date of Patent: December 10, 2019Assignee: PRATT & WHITNEY CANADA CORP.Inventors: Remy Synnott, Mohammed Ennacer, Chris Pater, Denis Blouin, Kapila Jain, Farough Mohammadi
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Patent number: 10502075Abstract: A component for a gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, a platform having a first path side and a second path side and a platform cooling circuit disposed on one of the first path side and the second path side of the platform. The platform cooling circuit includes a first core cavity, a cavity in fluid communication with the first core cavity, and a cover plate positioned to cover at least the cavity.Type: GrantFiled: November 24, 2015Date of Patent: December 10, 2019Assignee: United Technologies CorporationInventors: Michael Leslie Clyde Papple, Russell J. Bergman, Mohammed Ennacer, Shawn J. Gregg, Dominic J. Mongillo
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Patent number: 10443434Abstract: A turbine airfoil segment includes inner and outer platforms that are joined by at least one airfoil. The airfoil includes leading and trailing edges that are joined by spaced apart first and second sides to provide an exterior airfoil surface. At least one of the inner and outer platforms includes film cooling holes that have external breakout points that are located in substantial conformance with the Cartesian coordinates set forth in Table 1 for the inner platform or Table 2 for the outer platform. The Cartesian coordinates are provided by an axial coordinate, a circumferential coordinate, and a radial coordinate, relative to a zero-coordinate. The film cooling holes have a diametrical surface tolerance relative to the specified coordinates of 0.20 inches (5.0 mm).Type: GrantFiled: November 30, 2015Date of Patent: October 15, 2019Assignee: UNITED TECHNOLOGIES CORPORATIONInventors: Mohammed Ennacer, Russell J. Bergman, Jason B. Moran
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Publication number: 20190178101Abstract: A turbine shroud segment has a body having a radially outer surface and a radially inner surface extending axially between a leading edge and a trailing edge and circumferentially between a first and a second lateral edge. A first serpentine channel is disposed axially along the first lateral edge. A second serpentine channel is disposed axially along the second lateral edge. The first and second serpentine channels each define a tortuous path including axially extending passages between a front inlet proximate the leading edge and a rear outlet at the trailing edge of the body.Type: ApplicationFiled: December 13, 2017Publication date: June 13, 2019Inventors: Remy SYNNOTT, Mohammed ENNACER, Chris PATER, Denis BLOUIN, Kapila JAIN, Farough MOHAMMADI
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Publication number: 20190178102Abstract: A turbine shroud segment has a body extending axially between a leading edge and a trailing edge and circumferentially between a first and a second lateral edge. A core cavity is defined in the body and extends axially from a front end adjacent the leading edge to a rear end adjacent to the trailing edge. A plurality of cooling inlets and outlets are respectively provided along the front end and the rear end of the core cavity. A crossover wall extends across the core cavity and defines a row of crossover holes configured to accelerate the flow of coolant directed into the core cavity via the cooling inlets. The crossover wall is positioned to accelerate the coolant flow at the beginning of the cooling scheme where the shroud segment is the most thermally solicited.Type: ApplicationFiled: December 13, 2017Publication date: June 13, 2019Inventors: Remy SYNNOTT, Mohammed ENNACER, Chris PATER, Denis BLOUIN, Kapila JAIN, Farough MOHAMMADI