Patents by Inventor Ronald Scott Bunker

Ronald Scott Bunker 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).

  • Riblets For A Flowpath Surface Of A Turbomachine
    Publication number: 20170234134
    Abstract: A component stage for a turbomachine includes a component section. The component section includes a flowpath surface at least partially exposed to a core air flowpath defined by the turbomachine, when the component stage is installed in the turbomachine. The component further includes a plurality of sequentially arranged riblets on the flowpath surface, the plurality of sequentially arranged riblets customized for an anticipated location of the flowpath surface within the turbomachine by defining one or both of a non-uniform geometry or a non-uniform spacing.
    Type: Application
    Filed: February 12, 2016
    Publication date: August 17, 2017
    Inventor: Ronald Scott Bunker
  • IMPINGEMENT HOLES FOR A TURBINE ENGINE COMPONENT
    Publication number: 20170234139
    Abstract: An apparatus for an impingement hole for an engine component of a gas turbine engine includes an impingement baffle. The impingement baffle is spaced from an impingement surface and includes a plurality of impingement holes for providing an impingement flow to the impingement surface. The impingement holes can have an angled upstream edge such that an inlet has a greater cross-sectional area than an outlet. The walls of the impingement holes can have a hood to provide a higher shear flow content to minimize dust accumulation on the impingement surface.
    Type: Application
    Filed: February 13, 2016
    Publication date: August 17, 2017
    Inventors: Ronald Scott Bunker, Jason Randolph Allen
  • AIRFOIL HAVING IMPINGEMENT OPENINGS
    Publication number: 20170234146
    Abstract: An airfoil for a turbine engine having a perimeter wall bounding an interior and defining a pressure side and a suction side, a radially extending rib located within the interior and spaced from the leading edge to define a radially extending leading edge chamber, and at least one impingement opening in the rib defining a flow path aligned with the leading edge.
    Type: Application
    Filed: February 16, 2016
    Publication date: August 17, 2017
    Inventors: Ronald Scott Bunker, Robert David Briggs, Timothy Deryck Stone
  • Surface Contouring
    Publication number: 20170234537
    Abstract: A component for at least one of a combustion section or a turbine section of a gas turbine engine is provided. The combustion section and turbine section of the gas turbine engine at least partially define a core air flowpath, and the component includes a wall. The wall, in turn, includes a hot side and an opposite cold side. The hot side is exposed to and at least partially defines the core air flowpath when the component is installed in the gas turbine engine. The wall is manufactured to include surface contouring on the cold side of the wall to structurally accommodate a thermal management feature of the wall.
    Type: Application
    Filed: February 12, 2016
    Publication date: August 17, 2017
    Inventor: Ronald Scott Bunker
  • AIRFOIL FOR A GAS TURBINE ENGINE
    Publication number: 20170234140
    Abstract: A method and apparatus for an airfoil in a gas turbine engine can include an outer surface bounding an interior. At least one flow channel can be defined among one or more full-length and partial-length ribs to further define a cooling circuit within the airfoil. The cooling circuit can have at least one tip turn at the partial-length rib, having at least on fastback turbulator disposed at least partially within the tip turn.
    Type: Application
    Filed: February 13, 2016
    Publication date: August 17, 2017
    Inventors: Ronald Scott Bunker, Robert David Briggs, Jason Randolph Allen, Kevin Robert Feldmann
  • Airfoil Trailing Edge Cooling
    Publication number: 20170234138
    Abstract: A turbine airfoil for a gas turbine engine includes a pressure sidewall extending along a spanwise direction, and from a leading edge of the airfoil towards the trailing edge of the airfoil. The turbine airfoil additionally includes a suction sidewall also extending along the spanwise direction, and from the leading edge towards the trailing edge. The pressure sidewall and suction sidewall define a cooling air cavity therebetween, and one or both of the pressure sidewall and suction sidewall define a trailing edge cooling channel extending from the cooling air cavity substantially to the trailing edge. Additionally, one or both of the pressure sidewall and suction sidewall include a plurality of pressure drop members extending partially into the trailing edge cooling channel for reducing an amount of cooling air flowing therethrough from the cooling air cavity.
    Type: Application
    Filed: February 12, 2016
    Publication date: August 17, 2017
    Inventor: Ronald Scott Bunker
  • Thermal Stress Relief Of A Component
    Publication number: 20170234150
    Abstract: A component for a gas turbine engine includes a hot side wall, a plurality of connection walls, and a cold side wall. The hot side wall is exposed to a core air flowpath defined by the gas turbine engine. The cold side wall is spaced from the hot side wall and rigidly connected to the hot side wall through the plurality of connection walls. The hot side wall, connection walls, and cold side wall together define a cooling air cavity. The cold side wall defines a thermal stress relief slot for at least partially accommodating a relative thermal expansion between the hot side wall and the cold side wall to reduce an amount of thermal stress within the component during operation of the component.
    Type: Application
    Filed: February 12, 2016
    Publication date: August 17, 2017
    Inventor: Ronald Scott Bunker
  • Flowpath Contouring
    Publication number: 20170234161
    Abstract: A gas turbine engine includes a compressor section, combustion section, and turbine section. The turbine section includes a turbine component stage, the turbine component stage including a plurality of turbine components together including a flowpath surface along a circumferential direction of the gas turbine engine. The flowpath surface defines in part a core air flowpath of the gas turbine engine and further defines a contour along the circumferential direction. The contour repeats less frequently than once per turbine component to accommodate a hot gas streak through the turbine section.
    Type: Application
    Filed: February 12, 2016
    Publication date: August 17, 2017
    Inventor: Ronald Scott Bunker
  • COMPONENT COOLING FOR A GAS TURBINE ENGINE
    Publication number: 20170234225
    Abstract: An apparatus and method for an engine component having a hot surface adjacent a hot combustion gas flow and a cooling surface adjacent a cooling fluid flow can include at least one dimple for enhancing the cooling along the cooling surface. The dimple can be shaped having a head and a tail with the head disposed upstream of the tail to provide for reduced dust collection along the engine component.
    Type: Application
    Filed: February 13, 2016
    Publication date: August 17, 2017
    Inventors: Ronald Scott Bunker, Jeremy Clyde Bailey
  • Cooled airfoil trailing edge and method of cooling the airfoil trailing edge
    Patent number: 9732617
    Abstract: An airfoil and method of cooling a airfoil including a leading edge, a trailing edge, a suction side, a pressure side and at least one internal cooling channel configured to convey a cooling fluid, is provided. A plurality of trailing edge bleed slots are in fluid communication with the at least one internal cooling channel, wherein a downstream edge of the pressure side of the airfoil lies upstream of a downstream edge of the suction side to expose the plurality of trailing edge bleed slots proximate to the trailing edge of the airfoil. The at least one internal cooling channel is configured to supply the cooling fluid from a source of cooling fluid towards the plurality of trailing edge bleed slots. A plurality of obstruction features are disposed within the at least one internal cooling channel and at a downstream edge of the remaining pressure side.
    Type: Grant
    Filed: November 26, 2013
    Date of Patent: August 15, 2017
    Assignee: General Electric Company
    Inventors: Guanghua Wang, Ronald Scott Bunker, Jeremy Clyde Bailey, James Peter DeLancey, David Benjamin Helmer
  • GAS TURBINE ENGINE WITH A COOLING FLUID PATH
    Publication number: 20170226882
    Abstract: An apparatus and method of a gas turbine engine comprising a rotor having at least one disk with a rotor defining an axial face and a stator having at least one ring with a stator axial face confronting the rotor axial face, with terminal portions of the axial faces forming a fluid outlet there between. A recess formed in one of the axial faces defines a buffer cavity into which a wing extends from the other of the axial faces and having a surface confronting the fluid outlet. A flow reverser is further provided within at least one of the surface or the terminal portion of the other of the axial faces.
    Type: Application
    Filed: February 5, 2016
    Publication date: August 10, 2017
    Inventors: Jonathan Russell Ratzlaff, Ronald Scott Bunker
  • ENGINE COMPONENT ASSEMBLY
    Publication number: 20170191417
    Abstract: An engine component assembly includes a first engine component having a hot surface in thermal communication with a hot combustion gas flow and a cooling surface, and a second engine component having a first surface in fluid communication with a cooling fluid flow and a second surface spaced from the cooling surface to define a space. A cooling aperture extends through the second engine component. A cooling feature extends from the cooling surface of the first engine component, and is oriented relative to the cooling aperture such that the cooling fluid flow is orthogonal and non-orthogonal to different portions of the cooling feature.
    Type: Application
    Filed: January 6, 2016
    Publication date: July 6, 2017
    Inventors: Ronald Scott Bunker, Robert Frederick Bergholz, Jason Randolph Allen, Robert David Briggs, Byron Andrew Pritchard, JR., Kevin Robert Feldmann, Curtis Walton Stover, Zachary Daniel Webster, Fernando Reiter, Michael Alan Meade
  • ENGINE COMPONENT WITH FILM COOLING
    Publication number: 20170167268
    Abstract: An engine component can comprise a cooled surface adjacent to a cooling flow and a hot surface adjacent to a hot flow of fluid. The component can comprise a wall separating the hot and cooling flows, defining the cooled surface and the hot surface, and having a plurality of film holes disposed in the wall. At least one turbulator and at least one film hole inlet can be disposed on the cooled surface. The turbulator and the inlet can be arranged to provide a steady flow of cooling fluid to the film hole. One arrangement can comprise spacing the film hole inlet at least two turbulator heights from the turbulator.
    Type: Application
    Filed: December 11, 2015
    Publication date: June 15, 2017
    Inventor: Ronald Scott Bunker
  • CLOSED LOOP COOLING METHOD FOR A GAS TURBINE ENGINE
    Publication number: 20170159675
    Abstract: An apparatus and method of cooling a gas turbine engine having a compressor with multiple, axially arranged stages of paired rotating blades and stationary vanes located between an outer compressor casing and inner compressor casing, comprising a closed loop cooling of the compressor by routing a liquid coolant through the vanes of at least some of the compressor stages and through an intercooler to draw heat into the liquid coolant and routing the heated liquid coolant through a heat exchanger comprising a heat exchanger.
    Type: Application
    Filed: December 3, 2015
    Publication date: June 8, 2017
    Inventors: Mohamed El Hacin Sennoun, Ronald Scott Bunker
  • TURBINE ENGINE WITH AN AIRFOIL HAVING A TIP SHELF OUTLET
    Publication number: 20170159451
    Abstract: An airfoil for a gas turbine engine includes a pressure side and a suction side, with a root and a tip wall. The pressure side and suction side extend beyond the tip wall to define a tip channel, defining a plurality of internal and external corners. A tip shelf can be defined in the pressure or suction sidewalls. A film hole can extend to the tip shelf, such that the film hole can be shaped to direct a flow of cooling fluid to the film hole to improve film cooling at the tip shelf.
    Type: Application
    Filed: May 24, 2016
    Publication date: June 8, 2017
    Inventors: Jared Peter Buhler, Ronald Scott Bunker, Victor Hugo Silva Correia, Brian Kenneth Corsetti
  • CLOSED LOOP COOLING METHOD AND SYSTEM WITH HEAT PIPES FOR A GAS TURBINE ENGINE
    Publication number: 20170159566
    Abstract: An apparatus and method of cooling a gas turbine engine including a core with a compressor section in which the compressor section includes a closed loop cooling circuit having a pump, at least one heat pipe extending from at least one of the stationary vanes, a heat exchanger located within the bypass air flow, and a coolant conduit passing fluidly coupled to the pump and heat exchanger and passing by the heat pipe. The pump pumps coolant through the coolant conduit to draw heat from the heat pipes into the coolant to form heated coolant, the heated coolant then passes through the heat exchanger, where the heat is rejected from the coolant to the bypass air to cool the coolant to form cooled coolant, which is then returned to the heat pipes.
    Type: Application
    Filed: December 3, 2015
    Publication date: June 8, 2017
    Inventors: Mohammed El Hacin Sennoun, Ronald Scott Bunker
  • GAS TURBINE ENGINE WITH FILLET FILM HOLES
    Publication number: 20170159449
    Abstract: An airfoil for a gas turbine engine can have an exterior wall and an interior wall, with each wall having a thickness. The walls can intersect to define a corner at the intersection. A cooling passage can be defined by the walls at or near the corner to provide fluid communication between the interior and exterior of the airfoil. A film hole can be disposed in the walls and can have a length and diameter to define a ratio of length to diameter, L/D. An arcuate fillet can be located in the corner to define an effective radius for the fillet. The effective radius can be at least 1.5 times larger than the thicknesses of the walls to provide for an increased length to diameter ratio for the film hole.
    Type: Application
    Filed: December 7, 2015
    Publication date: June 8, 2017
    Inventors: Jared Peter Buhler, Ronald Scott Bunker, Victor Hugo Silva Correia, Brian Kenneth Corsetti
  • INTERCOOLING SYSTEM AND METHOD FOR A GAS TURBINE ENGINE
    Publication number: 20170159568
    Abstract: A system and method of cooling a gas turbine engine having a compressor with multiple, axially arranged stages of paired rotating blades and stationary vanes located between an outer compressor casing and inner compressor casing, comprising a closed loop cooling of the compressor by extracting compressor air from a first location in the compressor, cooling the extracted air, and introducing the cooled, extracted air at a second location in the compressor, with the second location being upstream of the first location.
    Type: Application
    Filed: December 3, 2015
    Publication date: June 8, 2017
    Inventors: Mohamed El Hacin Sennoun, Ronald Scott Bunker
  • FILLET OPTIMIZATION FOR TURBINE AIRFOIL
    Publication number: 20170159450
    Abstract: A blade for a gas turbine engine comprises an airfoil having a pressure side and a suction side, with a root and a tip wall. The pressure side and suction side extend beyond the tip wall to define a tip channel, defining a plurality of internal and external corners. The corners comprise fillets to define a thickness being greater than the thickness for the pressure, suction, or tip walls. A film hole can extend through the fillet, such that the length of the film hole at the fillet can be increased to define an increased length-to-diameter ratio for the film hole to improve film cooling through the film hole.
    Type: Application
    Filed: December 7, 2015
    Publication date: June 8, 2017
    Inventors: Jared Peter Buhler, Ronald Scott Bunker, Victor Hugo Silva Correia, Brian Kenneth Corsetti
  • TRAILING EDGE COOLING FOR A TURBINE BLADE
    Publication number: 20170159452
    Abstract: A component for a gas turbine engine comprises an airfoil having an outer surface. One or more cooling passages can be disposed within the airfoil, having a cooling passage extending along a trailing edge. A plurality of cooling channels can extend from the cooling passage through the trailing edge. At least one flow element and at least one film hole can be disposed in the cooling channel or the trailing edge passage adjacent the cooling channel. The flow element and the film hole can be in a predetermined relationship with one another providing improved flow to the film hole.
    Type: Application
    Filed: December 3, 2015
    Publication date: June 8, 2017
    Inventors: Ronald Scott Bunker, Zachary Daniel Webster