Patents Examined by Lorne Meade
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Patent number: 10215092Abstract: A constant-volume combustion chamber for an aircraft turbine engine, including a compressed gas intake valve configured to adopt an open position and a closed position, and in the closed position blocking intake of compressed gas into the chamber, and a combusted gas exhaust valve configured to adopt a closed position, in the closed position blocking exhaust of combusted gas outside the chamber. At least one of the intake and exhaust valves includes at least one spherical plug.Type: GrantFiled: July 30, 2013Date of Patent: February 26, 2019Assignee: SAFRAN AIRCRAFT ENGINESInventor: Bernard Robic
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Patent number: 10208620Abstract: The invention relates to an overspeed protection device of an aircraft engine.Type: GrantFiled: July 21, 2015Date of Patent: February 19, 2019Assignees: SAFRAN ELECTRONICS & DEFENSE, SAFRAN HELICOPTER ENGINESInventors: Michael Montoya, Nicolas Marti, Stephen Langford, Rafael Samson
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Patent number: 10208626Abstract: A manifold mounting arrangement comprising: a discrete manifold module; a clevis arrangement having a clevis straddling a flange and pivotally secured to the flange by a clevis pin; and a locking bar configured to be coupled to the clevis and to constrain the manifold module relative to the clevis arrangement in the locality of the locking bar. The invention finds utility for mounting case cooling manifolds to turbine casing for a gas turbine engine.Type: GrantFiled: July 6, 2011Date of Patent: February 19, 2019Assignee: ROLLS-ROYCE plcInventors: Glyn Steel, John R. Howarth
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Patent number: 10174674Abstract: A device for extraction of bleed air from flowing air at or in an aircraft engine includes an adjusting device for adjustment of an inlet cross section of an opening for the bleed air in or at a wall during operation of the aircraft engine. The opening for the bleed air is arranged in or at a deformable base and the adjusting device acts on the deformable base for modifying the inlet cross section of the opening relative to the flowing air. The deformable base is part of a metallic housing in the aircraft engine.Type: GrantFiled: September 4, 2015Date of Patent: January 8, 2019Assignee: ROLLS-ROYCE DEUTSCHLAND LTD & CO KGInventor: Robert Thies
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Patent number: 10151271Abstract: A flow-management system may comprise a center body impermeable to air. A conical surface of the center body may face forward. A blocking surface of the center body may be coaxial with the conical surface and may comprise an annular recess. An annular ring may be aft of the center body and fluidly coupled with the blocking surface. A tube may encase the center body and annular ring. The annular ring may comprise an air-foil shape to direct a pulse to the blocking surface. The blocking surface may comprise a central peak and a circular ridge separated by the annular recess.Type: GrantFiled: June 29, 2015Date of Patent: December 11, 2018Assignee: UNITED TECHNOLOGIES CORPORATIONInventors: James D. Hill, Roger F. Blinn, Michael J. Cuozzo
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Patent number: 10145015Abstract: A method of and apparatus for producing electricity, hydrogen gas, oxygen gas, pure water using a geothermal heat are disclosed. A low voltage (such as less than 0.9V) is applied to a prepared solution containing hydrogen generating catalysts to generate hydrogen and oxygen. The hydrogen and oxygen are used to drive a gas turbine to generate electricity. The oxygen and hydrogen are combusted to generate heat and pure water. This process is advantageous in many aspects including desalinating salt/sea water using geothermal heat.Type: GrantFiled: December 3, 2013Date of Patent: December 4, 2018Assignee: Marine Power Products IncorporatedInventor: Jeffrey M. Carey
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Patent number: 10132498Abstract: Aspects of the disclosure are directed to a liner associated with a combustor of an aircraft engine, comprising: a thermal barrier coating, and a base metal, wherein the thermal barrier coating comprises a contoured surface on a flowpath side proximate to an exit of a hole formed by the thermal barrier coating and the base metal.Type: GrantFiled: January 20, 2015Date of Patent: November 20, 2018Assignee: United Technologies CorporationInventor: Steven W. Burd
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Patent number: 10132196Abstract: Gas turbine engine systems involving I-beam struts are provided. In this regard, a representative strut assembly for a gas turbine engine includes a first I-beam strut having first and second flanges spaced from each other and interconnected by a web, the first strut exhibiting a twist along a length of the web.Type: GrantFiled: June 29, 2012Date of Patent: November 20, 2018Assignee: United Technologies CorporationInventors: Joey Wong, Peter Chen, Dana P. Stewart, David N. Waxman, Michael A. Mike
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Patent number: 10107140Abstract: 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. The airfoil includes film cooling holes that have external breakout points that are located in substantial conformance with the Cartesian coordinates set forth in one of Tables 1 and 2. The Cartesian coordinates are provided by an axial coordinate, a circumferential coordinate, and a radial coordinate, relative to a zero-coordinate, and the film cooling holes have a diametrical surface tolerance relative to the specified coordinates of 0.20 inches (5.0 mm).Type: GrantFiled: December 3, 2015Date of Patent: October 23, 2018Assignee: UNITED TECHNOLOGIES CORPORATIONInventors: Mohammed Ennacer, Russell J. Bergman, Jason B. Moran
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Patent number: 10094566Abstract: A system having a gas turbine engine is provided. The gas turbine engine includes a turbine and a combustor coupled to the turbine. The combustor includes a combustion chamber, one or more fuel nozzles upstream from the combustion chamber, and a head end having an end cover assembly. The end cover assembly includes an oxidant inlet configured to receive an oxidant flow, a central oxidant passage, and at least one fuel supply passage. The central oxidant passage is in fluid communication with the oxidant inlet, and the central oxidant passage is configured to route the oxidant flow to the one or more fuel nozzles. The at least one fuel supply passage is configured to receive a fuel flow and route the fuel flow into the one or more fuel nozzles.Type: GrantFiled: February 2, 2016Date of Patent: October 9, 2018Assignees: General Electric Company, ExxonMobil Upstream Research CompanyInventors: Bradford David Borchert, Jesse Edwin Trout, Scott Robert Simmons, Almaz Valeev, Ilya Aleksandrovich Slobodyanskiy, Igor Petrovich Sidko, Leonid Yul'evich Ginesin
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Patent number: 10094567Abstract: A dual-fuel burning gas turbine combustor having a diffusive combustion burner to burn a liquid fuel and a gaseous fuel placed at the axis of the gas turbine combustor and a plurality of pre-mixing combustion burners to burn a liquid fuel and a gaseous fuel placed on an outer circumferential side of the diffusive combustion burner, each pre-mixing combustion burner having a liquid fuel nozzle, a plurality of gaseous fuel spray holes, a plurality of air holes, and a pre-mixing chamber to mix gaseous fuel and air, wherein each pre-mixing combustion burner has a double pipe sleeve at a connected portion between end cover and the pre-mixing combustion burner, and the double pipe sleeve has an inner sleeve having a gaseous fuel flow path, an outer sleeve positioned on an outer circumferential side of the inner sleeve, and a circular spacing formed between the inner sleeve and the outer sleeve.Type: GrantFiled: August 18, 2014Date of Patent: October 9, 2018Assignee: Mitsubishi Hitachi Power Systems, Ltd.Inventors: Shota Igarashi, Hirokazu Takahashi
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Patent number: 10077739Abstract: A cascade-variable area fan nozzle system (the “CVAFN System”) is provided. The CVAFN system may comprise a cascade portion and a variable area fan nozzle (“VAFN”) portion. The VAFN portion may include a VAFN panel. The cascade and VAFN panel may be integrally formed with one another. The CVAFN system may include an actuation system (e.g., a jack screw) that is configured to translate the cascade and VAFN panel forward and aft. The cascade and VAFN panel may be translated aft in response to activation of the thrust reverser and/or CVAFN system.Type: GrantFiled: April 24, 2014Date of Patent: September 18, 2018Assignee: ROHR, INC.Inventor: Norman J. James
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Patent number: 10072612Abstract: Provided herein are various improvements to rocket engine components and rocket engine operational techniques. In one example, a rocket engine propellant injection apparatus is provided that includes a manifold formed into a single body by an additive manufacturing process and comprising a fuel cavity and an oxidizer cavity. The manifold also includes one or more propellant feed stubs, the one or more propellant feed stubs protruding from the manifold and formed into the single body of the manifold by the additive manufacturing process, with at least a first stub configured to carry fuel to the fuel cavity and at least a second stub configured to carry oxidizer to the oxidizer cavity. The manifold also includes a plurality of injection features formed by apertures in a face of the manifold, ones of the plurality of injection features configured to inject the fuel and the oxidizer for combustion.Type: GrantFiled: October 4, 2016Date of Patent: September 11, 2018Assignee: Vector Launch Inc.Inventors: Christopher Bostwick, John Garvey, Christopher Anderson, Eric Besnard
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Patent number: 10072610Abstract: The invention relates to a feed device for feeding a thrust chamber (10) of a rocket engine (100) with first and second propellants. According to the invention, a first feed circuit (16) of the thrust chamber (10) comprises a turbopump (22) having at least one pump (22a) for pumping the first propellant from a first tank (12), and a turbine (22b) mechanically coupled to said pump (22a). The first feed circuit connects an outlet of the pump to an inlet of the turbine via a heat exchanger (24) configured to heat the first propellant with heat generated by the thrust chamber, in order to actuate the turbine. According to the invention, a second feed circuit (18) is configured to feed the thrust chamber with second propellant from a second tank (14) that is configured to be pressurized. The invention also provides a method of feeding a rocket engine thrust chamber with first and second propellants.Type: GrantFiled: November 4, 2013Date of Patent: September 11, 2018Assignee: Arianegroup SASInventors: Jean Michel Sannino, François Lassoudiere, David Hayoun
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Patent number: 10066581Abstract: The invention relates to an aeroengine after-body assembly comprising an exhaust casing made of metal having a plurality of arms extending radially between an inner shroud and an outer shroud. The assembly comprises at least one axisymmetric part made of composite material extending between an upstream end fastened to said exhaust casing and a downstream end that is free. In accordance with the invention, the axisymmetric part has an annular portion at its upstream end, which annular portion includes a plurality of slots defining between them a plurality of resilient fastener tabs. Each slot co-operates with an arm of the exhaust casing, which further includes fastener parts attached to the resilient fastening tabs.Type: GrantFiled: March 28, 2013Date of Patent: September 4, 2018Assignee: SAFRAN NACELLESInventors: Gautier Mecuson, Eric Conete, Benoit Carrere, Eric Philippe
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Patent number: 10066836Abstract: Gas turbine engine systems and methods involving enhanced fuel dispersion are provided. In this regard, a representative method for operating a gas turbine engine includes: providing a gas path through the engine; introducing a spray of fuel along the gas path downstream of a turbine of the engine; and impinging the spray of fuel with a relatively higher velocity flow of air such that atomization of the fuel is increased.Type: GrantFiled: July 16, 2015Date of Patent: September 4, 2018Assignee: United Technologies CorporationInventors: Timothy S. Snyder, Steven W. Burd, Randal G. McKinney, George F. Titterton, III
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Patent number: 10060356Abstract: A rotational machine such as a turbocompressor has a fluid recovery system for recovering leaked working fluid such as gaseous helium in a helium circuit which has leaked past a shaft seal, a purifier being provided for removing contaminants from the working fluid, and turbocompressor may have one fluid such as helium or hydrogen working through one turbo component such as a turbine thereof and a second working fluid such as air or helium working through a second turbo component such as a compressor thereof, the rotational machine being installable in an engine of a flying machine.Type: GrantFiled: June 5, 2014Date of Patent: August 28, 2018Assignee: Reaction Engines LtdInventor: Alan Bond
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Patent number: 10060629Abstract: A combustor is provided. The combustor may comprise an axial fuel delivery system, and a radial fuel delivery system aft of the axial fuel delivery system. The radial fuel delivery system may be configured to direct fuel at least partially towards the axial fuel delivery system. A radial fuel delivery system is also provided. The system may comprise a combustor including a combustor liner, a mixer coupled to the combustor liner, and a nozzle disposed within the mixer, wherein the mixer and the nozzle are configured to direct fuel in a direction at least partially forward.Type: GrantFiled: February 20, 2015Date of Patent: August 28, 2018Assignee: UNITED TECHNOLOGIES CORPORATIONInventors: Wookyung Kim, Zhongtao Dai, Kristin Kopp-Vaughan
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Patent number: 10040563Abstract: An aircraft engine variable area fan nozzle structure disposed abaft a thrust reverser, including a sleeve translatable over a cascade array, comprises two semi cylindrical segments that can be axially translated and radially tilted to enlarge the fan duct exhaust area in order to optimize exhaust pressure and associated noise in high thrust circumstances such as on take-off, and to constrict that area under lower thrust conditions such as cruise. A pair of angularly adjacent segments can be moved by an actuator anchored to the fixed engine framework and independently of the thrust reverser translating sleeve. The tilting movement is imposed by the pivoting links of each segment to carriages that ride in a non-linear trackway secured to a thrust reverser translating sleeve. The actuator can be of a dual concentric type which can independently drive the trust reverser and nozzle segments.Type: GrantFiled: September 25, 2014Date of Patent: August 7, 2018Inventor: Geoffrey P. Pinto
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Patent number: 10041676Abstract: A combustor assembly for a gas turbine engine defining an axial direction is provided. The combustor assembly generally includes a liner, an annular dome, and a deflector plate. The liner at least partially defines a combustion chamber. The annular dome defines a first cavity and has a plurality of impingement holes, and the deflector plate defines a conical surface and a flat surface, wherein a plurality of cooling holes are defined through the deflector plate. The annular dome and the deflector plate are positioned together to define a second cavity that is in fluid communication with the first cavity through the plurality of impingement holes. In addition, the second cavity is in fluid communication with the combustion chamber through the plurality of cooling holes.Type: GrantFiled: July 8, 2015Date of Patent: August 7, 2018Assignee: General Electric CompanyInventors: Nayan Vinodbhai Patel, Kwanwoo Kim, Chad Holden Sutton, Duane Douglas Thomsen, Craig Alan Gonyou, Harris Daniel Abramson, James A. Russo, Robert Andrew Stowers, Shanwu Wang, Anquan Wang