Patents by Inventor Simon W. Evans
Simon W. Evans 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).
-
Patent number: 11773866Abstract: A compressor section for a gas turbine engine includes a blade including a platform, a tip and an airfoil extending between the platform and tip. The airfoil includes a root portion adjacent to the platform, a midspan portion and a tip portion. Each of the root portion, midspan portion and tip portion define a meridional velocity at stage exit with the tip portion including a first meridional velocity greater than a second meridional velocity of the midspan portion. A blade for an axial compressor of a gas turbine engine and a method of operating a compressor section of a gas turbine engine are also disclosed.Type: GrantFiled: February 1, 2022Date of Patent: October 3, 2023Assignee: RTX CORPORATIONInventors: Simon W. Evans, Yuan Dong, Sean Nolan
-
Publication number: 20230193855Abstract: An airfoil for a gas turbine engine including an airfoil body extending between a leading edge and a trailing edge and between a pressure side and a suction side. The airfoil body includes a strut portion extending from the leading edge and a flap portion extending from the trailing edge. The flap portion is pivotable relative to the strut portion. A flexible skin surrounds both the strut portion and the flap portion on both the pressure side and the suction side.Type: ApplicationFiled: February 23, 2023Publication date: June 22, 2023Inventors: Michael M. Joly, Zaffir A. Chaudhry, Andrzej Ernest Kuczek, Om P. Sharma, Georgi Kalitzin, Simon W. Evans
-
Patent number: 11655778Abstract: A gas turbine engine includes a fan section, a compressor section, and a turbine section. The fan section has a plurality of vane assemblies spaced circumferentially about an engine axis. The vane assemblies each include an airfoil extending between a leading edge and a trailing edge, a control rod extending through the airfoil, and a mechanism driven by the control rod to change the shape of the airfoil. A vane system for a gas turbine engine is also disclosed.Type: GrantFiled: August 6, 2021Date of Patent: May 23, 2023Assignee: Raytheon Technologies CorporationInventors: Michael M. Joly, Zaffir A. Chaudhry, Andrzej Ernest Kuczek, Om P. Sharma, Georgi Kalitzin, Simon W. Evans
-
Publication number: 20230044195Abstract: A gas turbine engine includes a fan section, a compressor section, and a turbine section. The fan section has a plurality of vane assemblies spaced circumferentially about an engine axis. The vane assemblies each include an airfoil extending between a leading edge and a trailing edge, a control rod extending through the airfoil, and a mechanism driven by the control rod to change the shape of the airfoil. A vane system for a gas turbine engine is also disclosed.Type: ApplicationFiled: August 6, 2021Publication date: February 9, 2023Inventors: Michael M. Joly, Zaffir A. Chaudhry, Andrzej Ernest Kuczek, Om P. Sharma, Georgi Kalitzin, Simon W. Evans
-
Publication number: 20220154728Abstract: A compressor section for a gas turbine engine includes a blade including a platform, a tip and an airfoil extending between the platform and tip. The airfoil includes a root portion adjacent to the platform, a midspan portion and a tip portion. Each of the root portion, midspan portion and tip portion define a meridional velocity at stage exit with the tip portion including a first meridional velocity greater than a second meridional velocity of the midspan portion. A blade for an axial compressor of a gas turbine engine and a method of operating a compressor section of a gas turbine engine are also disclosed.Type: ApplicationFiled: February 1, 2022Publication date: May 19, 2022Inventors: Simon W. Evans, Yuan Dong, Sean Nolan
-
Patent number: 11248622Abstract: A compressor section for a gas turbine engine includes a blade including a platform, a tip and an airfoil extending between the platform and tip. The airfoil includes a root portion adjacent to the platform, a midspan portion and a tip portion. Each of the root portion, midspan portion and tip portion define a meridional velocity at stage exit with the tip portion including a first meridional velocity greater than a second meridional velocity of the midspan portion. A blade for an axial compressor of a gas turbine engine and a method of operating a compressor section of a gas turbine engine are also disclosed.Type: GrantFiled: September 2, 2016Date of Patent: February 15, 2022Assignee: Raytheon Technologies CorporationInventors: Simon W. Evans, Yuan Dong, Sean Nolan
-
Patent number: 11111811Abstract: A stator vane for a gas turbine engine section includes a stator vane having an airfoil extending between a leading edge and a trailing edge. The airfoil has a suction side and a pressure side. There is at least one piezoelectric actuator for changing a shape of at least one of the leading edge and the trailing edge. A gas turbine engine is also disclosed.Type: GrantFiled: July 2, 2019Date of Patent: September 7, 2021Assignee: Raytheon Technologies CorporationInventors: Michael M. Joly, Zaffir A. Chaudhry, Simon W. Evans, Gorazd Medic, Dilip Prasad
-
Patent number: 11078805Abstract: A compressor for use in a gas turbine engine including: an airfoil configured to rotate about an engine central longitudinal axis of the gas turbine engine; a casing, the casing including a radially inward surface; and a groove located within the casing opposite the airfoil, the groove is recessed in a radially outward direction from the radially inward surface of the casing, wherein the groove is defined by a forward groove wall, an aft groove wall opposite the forward groove wall, and a base groove interposed between the forward groove wall and the aft groove wall, and wherein the forward groove wall is operably shaped to generate an aft directed jet.Type: GrantFiled: April 15, 2019Date of Patent: August 3, 2021Assignee: RAYTHEON TECHNOLOGIES CORPORATIONInventors: Michael M. Joly, Simon W. Evans, Dilip Prasad, Gorazd Medic, Georgi Kalitzin, Dmytro Mykolayovych Voytovych
-
Patent number: 10947851Abstract: A rotor blade of a gas turbine engine includes a pressure side, and a suction side opposite the pressure side and defining a rotor blade profile therebetween, the pressure side and the suction side each extending from a blade root to a blade tip. The rotor blade defines a cross-sectional median line midway between the pressure side and the suction side. The cross-sectional median line extends in a generally radial direction from the blade root to a lean point between the blade root and the blade tip. The cross-sectional median line extends off of radial from the lean point to the blade tip, defining a lean of the rotor blade between the lean point and the blade tip.Type: GrantFiled: December 19, 2018Date of Patent: March 16, 2021Assignee: RAYTHEON TECHNOLOGIES CORPORATIONInventor: Simon W. Evans
-
Publication number: 20210003030Abstract: A stator vane for a gas turbine engine section includes a stator vane having an airfoil extending between a leading edge and a trailing edge. The airfoil has a suction side and a pressure side. There is at least one piezoelectric actuator for changing a shape of at least one of the leading edge and the trailing edge. A gas turbine engine is also disclosed.Type: ApplicationFiled: July 2, 2019Publication date: January 7, 2021Inventors: Michael M. Joly, Zaffir A. Chaudhry, Simon W. Evans, Gorazd Medic, Dilip Prasad
-
Publication number: 20200325790Abstract: A compressor for use in a gas turbine engine including: an airfoil configured to rotate about an engine central longitudinal axis of the gas turbine engine; a casing, the casing including a radially inward surface; and a groove located within the casing opposite the airfoil, the groove is recessed in a radially outward direction from the radially inward surface of the casing, wherein the groove is defined by a forward groove wall, an aft groove wall opposite the forward groove wall, and a base groove interposed between the forward groove wall and the aft groove wall, and wherein the forward groove wall is operably shaped to generate an aft directed jet.Type: ApplicationFiled: April 15, 2019Publication date: October 15, 2020Inventors: Michael M. Joly, Simon W. Evans, Dilip Prasad, Gorazd Medic, Georgi Kalitzin, Dmytro Mykolayovych Voytovych
-
Publication number: 20200200014Abstract: A rotor blade of a gas turbine engine includes a pressure side, and a suction side opposite the pressure side and defining a rotor blade profile therebetween, the pressure side and the suction side each extending from a blade root to a blade tip. The rotor blade defines a cross-sectional median line midway between the pressure side and the suction side. The cross-sectional median line extends in a generally radial direction from the blade root to a lean point between the blade root and the blade tip. The cross-sectional median line extends off of radial from the lean point to the blade tip, defining a lean of the rotor blade between the lean point and the blade tip.Type: ApplicationFiled: December 19, 2018Publication date: June 25, 2020Inventor: Simon W. Evans
-
Patent number: 10605260Abstract: Rotor of a gas turbine engines having a rotor hub and a plurality of blades extending from the rotor hub, wherein each blade has a full-span forward sweep along a leading edge of the blade that starts at an airfoil root of the blade at the hub and extends to a blade tip, wherein a sweep of a blade is a percentage of a root axial chord length of the respective blade.Type: GrantFiled: September 9, 2016Date of Patent: March 31, 2020Assignee: UNITED TECHNOLOGIES CORPORATIONInventors: Li Xing Pan, Yuan Dong, Simon W. Evans, Robert W. Fessenden, Sue-Li Kingsley Chuang, Sean Nolan
-
Publication number: 20180073517Abstract: Rotor of a gas turbine engines having a rotor hub and a plurality of blades extending from the rotor hub, wherein each blade has a full-span forward sweep along a leading edge of the blade that starts at an airfoil root of the blade at the hub and extends to a blade tip, wherein a sweep of a blade is a percentage of a root axial chord length of the respective blade.Type: ApplicationFiled: September 9, 2016Publication date: March 15, 2018Inventors: Li Xing Pan, Yuan Dong, Simon W. Evans, Robert W. Fessenden, Sue-Li Kingsley Chuang, Sean Nolan
-
Publication number: 20180066673Abstract: A compressor section for a gas turbine engine includes a blade including a platform, a tip and an airfoil extending between the platform and tip. The airfoil includes a root portion adjacent to the platform, a midspan portion and a tip portion. Each of the root portion, midspan portion and tip portion define a meridional velocity at stage exit with the tip portion including a first meridional velocity greater than a second meridional velocity of the midspan portion. A blade for an axial compressor of a gas turbine engine and a method of operating a compressor section of a gas turbine engine are also disclosed.Type: ApplicationFiled: September 2, 2016Publication date: March 8, 2018Inventors: Simon W. Evans, Yuan Dong, Sean Nolan