Patents by Inventor Craig William Higgins

Craig William Higgins 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).

  • Publication number: 20250172110
    Abstract: A gas turbine engine is provided. The gas turbine engine includes a turbomachine defining an engine inlet to an inlet duct, a fan duct inlet to a fan duct, and a core inlet to a core duct; a primary fan driven by the turbomachine; and a secondary fan located downstream of the primary fan within the inlet duct. The gas turbine engine defines a thrust to power airflow ratio between 3.5 and 100 and a core bypass ratio between 0.1 and 10, wherein the thrust to power airflow ratio is a ratio of an airflow through a bypass passage over the turbomachine plus an airflow through the fan duct to an airflow through the core duct, and wherein the core bypass ratio is a ratio of the airflow through the fan duct to the airflow through the core duct.
    Type: Application
    Filed: January 31, 2025
    Publication date: May 29, 2025
    Inventors: William Joseph Bowden, David Marion Ostdiek, Ian McKnight, Syed J. Khalid, Brandon Wayne Miller, Randy M. Vondrell, Craig William Higgins, Alexander Kimberley Simpson
  • Publication number: 20250146454
    Abstract: A gas turbine engine includes: a turbomachine comprising a compressor section, a combustion section, and a turbine section arranged in serial flow order, the turbomachine defining an engine inlet to an inlet duct, a fan duct inlet to a fan duct, and a core inlet to a core duct; a primary fan driven by the turbomachine; and a secondary fan located downstream of the primary fan within the inlet duct, the gas turbine engine defining a thrust to power airflow ratio between 3.5 and 100 and a core bypass ratio between 0.1 and 10, wherein the thrust to power airflow ratio is a ratio of an airflow through a bypass passage over the turbomachine plus an airflow through the fan duct to an airflow through the core duct, wherein the core bypass ratio is a ratio of the airflow through the fan duct to the airflow through the core duct.
    Type: Application
    Filed: January 10, 2025
    Publication date: May 8, 2025
    Inventors: Brandon Wayne Miller, Randy M. Vondrell, David Marion Ostdiek, Craig William Higgins, Alexander Kimberley Simpson, Syed Arif Khalid
  • Publication number: 20250146453
    Abstract: A gas turbine engine is provided. The gas turbine engine includes a turbomachine defining an engine inlet to an inlet duct, a fan duct inlet to a fan duct, and a core inlet to a core duct; a primary fan driven by the turbomachine; and a secondary fan located downstream of the primary fan within the inlet duct. The gas turbine engine defines a thrust to power airflow ratio between 3.5 and 100 and a core bypass ratio between 0.1 and 10, wherein the thrust to power airflow ratio is a ratio of an airflow through a bypass passage over the turbomachine plus an airflow through the fan duct to an airflow through the core duct, and wherein the core bypass ratio is a ratio of the airflow through the fan duct to the airflow through the core duct.
    Type: Application
    Filed: January 10, 2025
    Publication date: May 8, 2025
    Inventors: Darek Tomasz Zatorski, David Marion Ostdiek, Mohamed Osama, William Joseph Solomon, Brandon Wayne Miller, Randy M. Vondrell, Craig William Higgins, Alexander Kimberley Simpson
  • Publication number: 20250129756
    Abstract: A gas turbine engine includes a turbomachine defining an engine inlet to an inlet duct, a fan duct inlet to a fan duct, and a core inlet to a core duct, a primary fan driven by the turbomachine, a secondary fan located downstream of the primary fan within the inlet duct, and a booster located downstream of the secondary fan and comprising a booster rotor blade, an inlet guide vane, and booster cowl, the booster cowl separating an upstream portion of the fan duct into an upper fan duct having an upper fan duct inlet and a lower fan duct having a lower fan duct inlet, the upper fan duct inlet and lower fan duct inlet collectively forming the fan duct inlet, the inlet guide vane located forward of the booster rotor blade.
    Type: Application
    Filed: December 10, 2024
    Publication date: April 24, 2025
    Inventors: Randy M. Vondrell, Alexander Kimberley Simpson, David Marion Ostdiek, Craig William Higgins
  • Publication number: 20250116243
    Abstract: A gas turbine engine includes a turbomachine having a compressor section, a combustion section, and a turbine section arranged in serial flow order. The turbomachine defines an engine inlet to an inlet duct, a fan duct inlet to a fan duct, and a core inlet to a core duct. The primary fan is driven by the turbomachine, and a secondary fan is located downstream of the primary fan within the inlet duct. One or more actuation devices operably associated with the fan duct, the one or more actuation devices actuable to increase or decrease an exit area of the fan duct.
    Type: Application
    Filed: December 16, 2024
    Publication date: April 10, 2025
    Inventors: William Joseph Bowden, Keith Edward James Blodgett, Mustafa Dindar, David Cerra, James Hamilton Grooms, Brandon Wayne Miller, Randy M. Vondrell, David Marion Ostdiek, Craig Williams Higgins, Alexander Kimberely Simpson
  • Publication number: 20250012235
    Abstract: A gas turbine engine is provided. The gas turbine engine includes a turbomachine defining an engine inlet to an inlet duct, a fan duct inlet to a fan duct, and a core inlet to a core duct; a primary fan driven by the turbomachine; and a secondary fan located downstream of the primary fan within the inlet duct. The gas turbine engine defines a thrust to power airflow ratio between 3.5 and 100 and a core bypass ratio between 0.1 and 10, wherein the thrust to power airflow ratio is a ratio of an airflow through a bypass passage over the turbomachine plus an airflow through the fan duct to an airflow through the core duct, and wherein the core bypass ratio is a ratio of the airflow through the fan duct to the airflow through the core duct.
    Type: Application
    Filed: September 18, 2024
    Publication date: January 9, 2025
    Inventors: Brandon Wayne Miller, Randy M. Vondrell, David Marion Ostdiek, Craig Williams Higgins, Alexander Kimberely Simpson
  • Publication number: 20250003375
    Abstract: A gas turbine engine includes a turbomachine defining an engine inlet to an inlet duct, a fan duct inlet to a fan duct, and a core inlet to a core duct, a primary fan driven by the turbomachine, a secondary fan located downstream of the primary fan within the inlet duct, a heat exchanger disposed in the fan duct, and a booster upstream of the heat exchanger, the booster including a booster cowl extending into the fan duct, the booster cowl separating an upstream portion of the fan duct into an upper fan duct having an upper fan duct inlet and a lower fan duct having a lower fan duct inlet, the upper fan duct inlet and lower fan duct inlet collectively forming the fan duct inlet.
    Type: Application
    Filed: June 12, 2024
    Publication date: January 2, 2025
    Inventors: Joseph George Rose, Thomas Ory Moniz, Tsuguji Nakano, Jeffrey S. Spruill, Brandon Wayne Miller, Randy M. Vondrell, David Marion Ostdiek, Craig William Higgins, Alexander Kimberley Simpson
  • Publication number: 20240318613
    Abstract: A gas turbine engine is provided. The gas turbine engine includes a turbomachine defining an engine inlet to an inlet duct, a fan duct inlet to a fan duct, and a core inlet to a core duct; a primary fan driven by the turbomachine; and a secondary fan located downstream of the primary fan within the inlet duct. The gas turbine engine defines a thrust to power airflow ratio between 3.5 and 100 and a core bypass ratio between 0.1 and 10, wherein the thrust to power airflow ratio is a ratio of an airflow through a bypass passage over the turbomachine plus an airflow through the fan duct to an airflow through the core duct, and wherein the core bypass ratio is a ratio of the airflow through the fan duct to the airflow through the core duct.
    Type: Application
    Filed: May 31, 2024
    Publication date: September 26, 2024
    Inventors: Brandon Wayne Miller, Randy M. Vondrell, David Marion Ostdiek, Craig Williams Higgins, Alexander Kimberley Simpson, Syed Arif Khalid, Jeffrey S. Spruill, Daniel Lawrence Tweedt, William Joseph Solomon, Kevin Edward Hinderliter
  • Publication number: 20240309829
    Abstract: A gas turbine engine is provided. The gas turbine engine includes a turbomachine defining an engine inlet to an inlet duct, a fan duct inlet to a fan duct, and a core inlet to a core duct; a primary fan driven by the turbomachine; and a secondary fan located downstream of the primary fan within the inlet duct. The gas turbine engine defines a thrust to power airflow ratio between 3.5 and 100 and a core bypass ratio between 0.1 and 10, wherein the thrust to power airflow ratio is a ratio of an airflow through a bypass passage over the turbomachine plus an airflow through the fan duct to an airflow through the core duct, and wherein the core bypass ratio is a ratio of the airflow through the fan duct to the airflow through the core duct.
    Type: Application
    Filed: May 28, 2024
    Publication date: September 19, 2024
    Inventors: Brandon Wayne Miller, Randy M. Vondrell, David Marion Ostdiek, Craig William Higgins, Alexander Kimberley Simpson
  • Patent number: 12031504
    Abstract: A gas turbine engine is provided. The gas turbine engine includes a turbomachine defining an engine inlet to an inlet duct, a fan duct inlet to a fan duct, and a core inlet to a core duct; a primary fan driven by the turbomachine; and a secondary fan located downstream of the primary fan within the inlet duct. The gas turbine engine defines a thrust to power airflow ratio between 3.5 and 100 and a core bypass ratio between 0.1 and 10, wherein the thrust to power airflow ratio is a ratio of an airflow through a bypass passage over the turbomachine plus an airflow through the fan duct to an airflow through the core duct, and wherein the core bypass ratio is a ratio of the airflow through the fan duct to the airflow through the core duct.
    Type: Grant
    Filed: August 2, 2022
    Date of Patent: July 9, 2024
    Assignee: General Electric Company
    Inventors: Brandon Wayne Miller, Randy M. Vondrell, David Marion Ostdiek, Craig William Higgins, Alexander Kimberley Simpson
  • Publication number: 20240044304
    Abstract: A gas turbine engine is provided. The gas turbine engine includes a turbomachine defining an engine inlet to an inlet duct, a fan duct inlet to a fan duct, and a core inlet to a core duct; a primary fan driven by the turbomachine; and a secondary fan located downstream of the primary fan within the inlet duct. The gas turbine engine defines a thrust to power airflow ratio between 3.5 and 100 and a core bypass ratio between 0.1 and 10, wherein the thrust to power airflow ratio is a ratio of an airflow through a bypass passage over the turbomachine plus an airflow through the fan duct to an airflow through the core duct, and wherein the core bypass ratio is a ratio of the airflow through the fan duct to the airflow through the core duct.
    Type: Application
    Filed: August 2, 2022
    Publication date: February 8, 2024
    Inventors: Brandon Wayne Miller, Randy M. Vondrell, David Marion Ostdiek, Craig William Higgins, Alexander Kimberley Simpson
  • Patent number: 11618580
    Abstract: A method for operating a hybrid electric propulsion system of an aircraft, the hybrid electric propulsion system comprising a turbomachine, an electric machine coupled to the turbomachine, and a propulsor coupled to the turbomachine, the method comprising: operating the turbomachine to drive the propulsor; receiving data indicative of a failure condition of the hybrid electric propulsion system; and extracting power from the turbomachine using the electric machine to slow down one or more rotating components of the turbomachine in response to receiving the data indicative of the failure condition.
    Type: Grant
    Filed: January 29, 2021
    Date of Patent: April 4, 2023
    Assignee: General Electric Company
    Inventors: Darek Tomasz Zatorski, Craig William Higgins
  • Publication number: 20220063825
    Abstract: A method for operating a hybrid electric propulsion system of an aircraft, the hybrid electric propulsion system comprising a turbomachine, an electric machine coupled to the turbomachine, and a propulsor coupled to the turbomachine, the method comprising: operating the turbomachine to drive the propulsor; receiving data indicative of a failure condition of the hybrid electric propulsion system; and extracting power from the turbomachine using the electric machine to slow down one or more rotating components of the turbomachine in response to receiving the data indicative of the failure condition.
    Type: Application
    Filed: January 29, 2021
    Publication date: March 3, 2022
    Inventors: Darek Tomasz Zatorski, Craig William Higgins
  • Patent number: 10968771
    Abstract: A method and icing effects mitigation system are provided. The icing effects mitigation system includes a fluid duct configured to channel a first flow of fluid through the fluid duct from a duct opening to a rotatable member at least partially positioned within the fluid duct. The rotatable member includes a radially inner rotatable portion and a radially outer rotatable portion. The icing effects mitigation system also includes a duct member extending through the fluid duct in a direction approximately orthogonal to a direction of the first flow of fluid. The duct member is configured to channel a second flow of a second fluid therethrough that causes ice accreted on the duct member to shed on a trajectory that impacts the rotatable member at the radially inner portion.
    Type: Grant
    Filed: January 12, 2017
    Date of Patent: April 6, 2021
    Assignee: GENERAL ELECTRIC COMPANY
    Inventors: David William Crall, Charles Daniel Califf, Craig William Higgins, Erich Alois Krammer
  • Publication number: 20180195408
    Abstract: A method and icing effects mitigation system are provided. The icing effects mitigation system includes a fluid duct configured to channel a first flow of fluid through the fluid duct from a duct opening to a rotatable member at least partially positioned within the fluid duct. The rotatable member includes a radially inner rotatable portion and a radially outer rotatable portion. The icing effects mitigation system also includes a duct member extending through the fluid duct in a direction approximately orthogonal to a direction of the first flow of fluid. The duct member is configured to channel a second flow of a second fluid therethrough that causes ice accreted on the duct member to shed on a trajectory that impacts the rotatable member at the radially inner portion.
    Type: Application
    Filed: January 12, 2017
    Publication date: July 12, 2018
    Inventors: David William Crall, Charles Daniel Califf, Craig William Higgins, Erich Alois Krammer
  • Patent number: 9581085
    Abstract: Embodiments of hot streak alignment for gas turbine durability include structures and methods to align hot streaks with the leading edges of aligned first stage nozzle vanes in order to improve mixing of the hot streaks with cooling air at a stator nozzle of a first stage turbine and reduce usage of cooling air at first stage non-aligned stator nozzle vanes disposed adjacent to the aligned stator vanes.
    Type: Grant
    Filed: March 15, 2013
    Date of Patent: February 28, 2017
    Assignee: General Electric Company
    Inventors: Robert Joseph Bartz, Joseph Steven Bubnick, Scott Matthew Bush, Kevin Robert Feldmann, Robert Alan Frederick, Craig William Higgins
  • Patent number: 9260974
    Abstract: Active clearance control ejector systems and methods are disclosed. An example system may include an air ejector; an entrance pipe arranged to supply compressor bleed air to the air ejector; an inlet duct arranged to supply fan bypass air to the air ejector; and a supply pipe arranged to receive ejector outlet air from the air ejector and supply the ejector outlet air to an active clearance control system. The ejector outlet air may include a mixture of the compressor bleed air and the fan bypass air. The air ejector may include a venturi arranged to conduct the compressor bleed air therethrough, thereby drawing the fan bypass air into the air ejector.
    Type: Grant
    Filed: September 25, 2012
    Date of Patent: February 16, 2016
    Assignee: General Electric Company
    Inventors: William Howard Hasting, David William Crall, Craig William Higgins, Erich Alois Krammer
  • Patent number: 8608436
    Abstract: A connection between an annular sleeve on a gas turbine engine first rotor component includes an annular tapered conical slot extending axially inwardly into the sleeve from an annular opening and tapers axially into the sleeve from the opening. An annular rim of a second rotor component is received within the conical slot and includes a rim inner conical surface mating with and contacting a sleeve inner conical surface in part bounding the conical slot. A sleeve cylindrical outer surface in part bounds the conical slot in the sleeve and a rim outer cylindrical surface on the annular rim contacts the sleeve cylindrical outer surface. An annular lip may extend radially inwardly from the sleeve and sleeve inner conical surface at the annular opening to the conical slot. The connection may be used between a first stage disk and an interstage seal.
    Type: Grant
    Filed: August 31, 2010
    Date of Patent: December 17, 2013
    Assignee: General Electric Company
    Inventors: Daniel Edward Wines, James Hamilton Grooms, Craig William Higgins
  • Patent number: 8172506
    Abstract: A method and system for a rotatable member of a turbine engine are provided. The rotatable member includes a substantially cylindrical shaft rotatable about a longitudinal axis, and a hub coupled to the cylindrical shaft through a conical shaft portion wherein the conical shaft portion includes a plurality of circumferentially-spaced air passages and wherein at least one of the plurality of air passages includes a non-circular cross section.
    Type: Grant
    Filed: November 26, 2008
    Date of Patent: May 8, 2012
    Assignee: General Electric Company
    Inventors: Rafal Piotr Pieczka, Craig William Higgins, Piotr Edward Kobek, Jan Franciszek Sikorski, Edward Donald Michaels
  • Publication number: 20120051917
    Abstract: A connection between an annular sleeve on a gas turbine engine first rotor component includes an annular tapered conical slot extending axially inwardly into the sleeve from an annular opening and tapers axially into the sleeve from the opening. An annular rim of a second rotor component is received within the conical slot and includes a rim inner conical surface mating with and contacting a sleeve inner conical surface in part bounding the conical slot. A sleeve cylindrical outer surface in part bounds the conical slot in the sleeve and a rim outer cylindrical surface on the annular rim contacts the sleeve cylindrical outer surface. An annular lip may extend radially inwardly from the sleeve and sleeve inner conical surface at the annular opening to the conical slot. The connection may be used between a first stage disk and an interstage seal.
    Type: Application
    Filed: August 31, 2010
    Publication date: March 1, 2012
    Inventors: Daniel Edward Wines, James Hamilton Grooms, Craig William Higgins