Patents by Inventor Richard Geoffrey STRETTON
Richard Geoffrey STRETTON 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: 10240536Abstract: A gas turbine engine is disclosed including a bifurcation fairing located in a bypass duct of the gas turbine engine and traversing the radial extent of the bypass duct. The bifurcation fairing has a delivery conduit inlet leading to a delivery conduit extending inside the bifurcation fairing, the delivery conduit being arranged in use for delivery of bypass air to one or more components of the gas turbine engine. A diverter conduit has a diverter conduit inlet from the delivery conduit upstream of the delivery conduit reaching the one or more components of the gas turbine engine. The diverter conduit has an outlet to a location other than the one or more components of the gas turbine engine.Type: GrantFiled: April 24, 2015Date of Patent: March 26, 2019Assignee: ROLLS-ROYCE plcInventors: Adam MacGregor Bagnall, Peter Beecroft, Richard Geoffrey Stretton, Philip Geoffrey Woodrow, Stephane Michel Marcel Baralon, Angus Roy Smith
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Patent number: 9970313Abstract: A unison ring assembly comprises a backing plate, and a plurality of pocket-forming portions. The backing plate is formed as an annular disc, with the annular disc being planar in a radial plane of the unison ring assembly. The plurality of pocket-forming portions are equi-spaced circumferentially around the axis of the unison ring assembly, and each of the plurality of pocket-forming portions is attached to the backing plate to form a corresponding radially extending pocket.Type: GrantFiled: May 19, 2015Date of Patent: May 15, 2018Assignee: ROLLS-ROYCE plcInventors: David Michael Beaven, Kenneth John Mackie, Richard Geoffrey Stretton
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Patent number: 9809317Abstract: An aircraft including a pylon attached to a gas turbine engine and a mounting system attaching the engine to the pylon. The mounting system including a first and a second frame each of three elongate members arranged in a triangle, each frame respectively arranged such that a core of the engine is positioned extending through an area defined between the three elongate members of each frame. Each frame forming at least part of a load bearing connection between the pylon and the engine. Each frame consisting of two portions, each portion corresponding to each side of the engine as attached to the pylon. The triangle formed by each frame being symmetrical about a plane separating the two portions. The engine is attached to the mounting system such that both frames are positioned axially forward of a radially extending projection of a first turbine stage in the core.Type: GrantFiled: November 17, 2014Date of Patent: November 7, 2017Assignee: ROLLS-ROYCE plcInventors: Richard Geoffrey Stretton, Kenneth Franklin Udall
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Patent number: 9771873Abstract: A gas turbine engine including an outlet guide vane and a bifurcation fairing is disclosed. The outlet guide vane is located in a bypass duct of the gas turbine engine downstream of a fan and is of aerofoil form. The bifurcation fairing traverses the radial extent of the bypass duct and has an upstream end that blends into a trailing edge of the outlet guide vane. The bifurcation fairing includes a scoop protruding outwards from its side corresponding to a pressure side of the upstream outlet guide vane. The scoop includes a forward facing inlet leading to a delivery conduit extending inside the bifurcation fairing for delivery in use of bypass air to one or more components of the gas turbine engine.Type: GrantFiled: May 6, 2015Date of Patent: September 26, 2017Assignee: ROLLS-ROYCE plcInventors: Peter Beecroft, Richard Geoffrey Stretton, Philip Geoffrey Woodrow, Stephane Michel Marcel Baralon, Angus Roy Smith
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Patent number: 9701412Abstract: An engine mount, for mounting the casing of an aircraft engine to the fuselage or wing of an aircraft, e.g. via a pylon, including first and second links having connector formations for connection to respective mounting and support formations on the engine casing and aircraft mounting structure. Each of the first and second links is typically connected by a pin between the mounting and support formations. Each link has a further connector formation arranged such that the connector formation of the first link is offset from the connector formation of the second link. The connector formations are joined by an intermediate link. The mount allows lateral forces and torque to be resolved in such a way as to substantially avoid lateral displacement of the engine away from a central plane. The intermediate link may also provide a failsafe catcher arrangement in the event that the first or second link fail.Type: GrantFiled: September 2, 2014Date of Patent: July 11, 2017Assignee: ROLLS-ROYCE plcInventors: Richard Geoffrey Stretton, Kenneth Franklin Udall
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Patent number: 9387923Abstract: A gas turbine engine (10) having an axial flow direction (X) therethrough in use. The gas turbine engine (10) comprises one or more rotor stages each comprising at least one rotor blade (120) having a root portion (122). The gas turbine engine (10) comprises a shroud (122) located upstream of one or more of the rotor stages relative to the axial flow direction (X). The shroud (122) defines a through passageway (128) extending between an inlet (130) and an outlet (132) which comprises a diffuser region (138). The diffuser region (138) is configured to reduce the axial velocity of air exiting the outlet (132) relative to air entering the diffuser portion (138) in use, wherein the outlet (132) is located such that air exiting the outlet (132) is directed substantially to the root portion (122) only of the rotor blades (120).Type: GrantFiled: June 11, 2013Date of Patent: July 12, 2016Assignee: ROLLS-ROYCE plcInventors: Richard Geoffrey Stretton, Nicholas Howarth
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Publication number: 20160167797Abstract: A mounting system is arranged in use for attaching a gas turbine engine to a pylon of an aircraft. The mounting system comprises a first frame of three elongate members arranged in a triangle. The first frame is arranged in use such that a core of the gas turbine engine is positioned extending through the area defined between the three elongate members. The first frame forms at least part of a load bearing connection between the pylon and gas turbine engine. The first frame consists of two portions, one corresponding to each side of the gas turbine engine when attached to the pylon. The triangle is symmetrical about a plane separating the two portions of the first frame.Type: ApplicationFiled: November 17, 2014Publication date: June 16, 2016Inventors: Richard Geoffrey STRETTON, Kenneth Franklin UDALL
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Patent number: 9346551Abstract: A fuselage mounted gas turbine engine installation the installation includes at least one propeller stage and a gas turbine core arranged in use to drive the propeller stage. The core is external to the fuselage and the rotational axes of the core and propeller stage are offset with respect to each other.Type: GrantFiled: February 5, 2014Date of Patent: May 24, 2016Assignee: ROLLS-ROYCE plcInventor: Richard Geoffrey Stretton
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Publication number: 20150361818Abstract: A unison ring assembly comprises a backing plate, and a plurality of pocket-forming portions. The backing plate is formed as an annular disc, with the annular disc being planar in a radial plane of the unison ring assembly. The plurality of pocket-forming portions are equi-spaced circumferentially around the axis of the unison ring assembly, and each of the plurality of pocket-forming portions is attached to the backing plate to form a corresponding radially extending pocket.Type: ApplicationFiled: May 19, 2015Publication date: December 17, 2015Inventors: David Michael BEAVEN, Kenneth John MACKIE, Richard Geoffrey STRETTON
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Publication number: 20150330309Abstract: A gas turbine engine is disclosed including a bifurcation fairing located in a bypass duct of the gas turbine engine and traversing the radial extent of the bypass duct. The bifurcation fairing has a delivery conduit inlet leading to a delivery conduit extending inside the bifurcation fairing, the delivery conduit being arranged in use for delivery of bypass air to one or more components of the gas turbine engine. A diverter conduit has a diverter conduit inlet from the delivery conduit upstream of the delivery conduit reaching the one or more components of the gas turbine engine. The diverter conduit has an outlet to a location other than the one or more components of the gas turbine engine.Type: ApplicationFiled: April 24, 2015Publication date: November 19, 2015Inventors: Adam MacGregor BAGNALL, Peter BEECROFT, Richard Geoffrey STRETTON, Philip Geoffrey WOODROW, Stephane Michel Marcel BARALON, Angus Roy SMITH
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Publication number: 20150330236Abstract: A gas turbine engine including an outlet guide vane and a bifurcation fairing is disclosed. The outlet guide vane is located in a bypass duct of the gas turbine engine downstream of a fan and is of aerofoil form. The bifurcation fairing traverses the radial extent of the bypass duct and has an upstream end that blends into a trailing edge of the outlet guide vane. The bifurcation fairing includes a scoop protruding outwards from its side corresponding to a pressure side of the upstream outlet guide vane. The scoop includes a forward facing inlet leading to a delivery conduit extending inside the bifurcation fairing for delivery in use of bypass air to one or more components of the gas turbine engine.Type: ApplicationFiled: May 6, 2015Publication date: November 19, 2015Inventors: Peter BEECROFT, Richard Geoffrey STRETTON, Philip Geoffrey WOODROW, Stephane Michel Marcel BARALON, Angus Roy SMITH
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Publication number: 20150300254Abstract: An air intake for an open rotor engine including a propulsive blade array having a plurality of blades each having a gas washed surface extending radially outwardly relative to an axis or rotation from a root end to a tip. Intake has first and second circumferential walls extending about axis of rotation at a location downstream of propulsive blade array. First and second walls are spaced in a radial direction to define an annular passage with opening having a height dimension extending a portion of the way along propulsive blade array span. Intake further includes an annular lip arranged about axis of rotation at a radial distance such that lip bifurcates the annular passage at a height which separates a boundary layer flow portion of intake flow from a remainder of intake flow. Lip may define a collection scroll for foreign object debris and cooling air for the engine.Type: ApplicationFiled: March 18, 2015Publication date: October 22, 2015Inventor: Richard Geoffrey STRETTON
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Publication number: 20150069176Abstract: An engine mount, for mounting the casing of an aircraft engine to the fuselage or wing of an aircraft, e.g. via a pylon, including first and second links having connector formations for connection to respective mounting and support formations on the engine casing and aircraft mounting structure. Each of the first and second links is typically connected by a pin between the mounting and support formations. Each link has a further connector formation arranged such that the connector formation of the first link is offset from the connector formation of the second link. The connector formations are joined by an intermediate link. The mount allows lateral forces and torque to be resolved in such a way as to substantially avoid lateral displacement of the engine away from a central plane. The intermediate link may also provide a failsafe catcher arrangement in the event that the first or second link fail.Type: ApplicationFiled: September 2, 2014Publication date: March 12, 2015Inventors: Richard Geoffrey STRETTON, Kenneth Franklin UDALL
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Publication number: 20140252159Abstract: A fuselage mounted gas turbine engine installation the installation includes at least one propeller stage and a gas turbine core arranged in use to drive the propeller stage. The core is external to the fuselage and the rotational axes of the core and propeller stage are offset with respect to each other.Type: ApplicationFiled: February 5, 2014Publication date: September 11, 2014Applicant: ROLLS-ROYCE PLCInventor: Richard Geoffrey STRETTON
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Publication number: 20140017067Abstract: A gas turbine engine (10) having an axial flow direction (X) therethrough in use. The gas turbine engine (10) comprises one or more rotor stages each comprising at least one rotor blade (120) having a root portion (122). The gas turbine engine (10) comprises a shroud (122) located upstream of one or more of the rotor stages relative to the axial flow direction (X). The shroud (122) defines a through passageway (128) extending between an inlet (130) and an outlet (132) which comprises a diffuser region (138). The diffuser region (138) is configured to reduce the axial velocity of air exiting the outlet (132) relative to air entering the diffuser portion (138) in use, wherein the outlet (132) is located such that air exiting the outlet (132) is directed substantially to the root portion (122) only of the rotor blades (120).Type: ApplicationFiled: June 11, 2013Publication date: January 16, 2014Inventors: Richard Geoffrey STRETTON, Nicholas HOWARTH