General Electric Company

Daniel Waslo

Mark Gregory Wotzak

Brandon Wayne Miller

Jeffrey Donald Clements

Aspi Rustom Wadia

Brian Lewis Devendorf

Joel Francis Kirk

Deepak Manohar Kamath

Kenneth Arthur Gould

Brian David Johnson

Michael Richardson-Bach

Martin Miles D'Angelo

The turbine includes a first turbine rotor disc, a second turbine rotor disc, a first source of cooling fluid, and a first cooling fluid conduit. The first turbine rotor disc has a first outer circumference and a first plurality of blades spaced along the first outer circumference. The second turbine rotor disc has a second outer circumference and a second plurality of blades spaced along the second outer circumference. The first cooling fluid conduit is configured to channel a first flow of cooling fluid from the first source of cooling fluid through the first turbine rotor disc to a blade of the second plurality of blades.

Turbine have a first and second rotor disc and a first and second cooling fluid conduit wherein the second cooling fluid conduit is extended through an annular axially extended bore having a radially outer extent defined by a radially innermost surface of the rotor discs

The present disclosure is directed to a gas turbine engine defining a radial direction, a circumferential direction, an axial centerline along a longitudinal direction, and an upstream end and a downstream end along the longitudinal direction. The gas turbine engine defines a core flowpath extended generally along the longitudinal direction. The gas turbine engine includes a first turbine rotor. The first turbine rotor includes an annular outer band disposed outward of the core flowpath along the radial direction. The first turbine rotor further includes a plurality of airfoils coupled to an inner diameter of the outer band in which the plurality of airfoils are extended generally inward along the radial direction. The outer band defines a plurality of airfoil cooling passages in which the plurality of airfoil cooling passages are extended at least partially in the radial direction in fluid communication with the plurality of airfoils.

Thermal structure for outer diameter mounted turbine blades

An apparatus and method for mounting a turbine engine to an aircraft can include an engine core for the turbine engine including a compressor section, a combustor section, and a turbine section in flow arrangement. At least one strut couples to the engine core about a single mount plane. A structural wall at least partially defining a mainstream flow path couples to the at least one strut and passes through the compressor section and the turbine section.

Turbine engine with single wall cantilevered architecture

A core engine includes a first tierod and a compressor rotor assembly including a plurality of compressor rotor disks arranged in a face to face orientation and spaced along the first tierod. The core engine includes a second tierod and a turbine rotor assembly including a plurality of turbine rotor disks arranged in a face to face orientation and spaced along the second tierod. The compressor rotor assembly is aft of the turbine rotor assembly.

Dual tierod assembly for a gas turbine engine and method of assembly thereof

The present disclosure is directed to a system for bleeding air from a compressed gas path of a gas turbine engine. The system includes an impeller positioned at a downstream end of a compressor in the gas turbine engine. The impeller includes an impeller hub, an impeller arm coupled to the impeller hub, and a plurality of circumferentially spaced apart impeller vanes extending radially outwardly from the impeller arm. The impeller arm defines an impeller arm aperture extending therethrough. A vortex spoiler is positioned radially inwardly from the impeller arm and defines a vortex spoiler passage extending radially therethrough. Bleed air flows from the compressed gas path radially inwardly through both the impeller arm aperture and the vortex spoiler passage.

Impeller-mounted vortex spoiler

A turbofan engine includes a compressor rotor, a plurality of main blades circumferentially mounted on the compressor rotor, and a plurality of spinner blades circumferentially mounted on the compressor rotor. Each of the spinner blades is located forward of a leading edge of a corresponding one of the main blades. Each of the spinner blades is circumferentially offset to the corresponding one of the main blades.

Turbofan engine's fan blade and setting method thereof

An in-line propeller gearbox of a turboprop gas turbine engine includes an epicyclic gearing arrangement that has a sun gear, a ring gear and at least one planet gear disposed between and meshing with both the sun gear and the ring gear. The propeller gearbox includes a disk brake that can be operated to slow down or stop altogether the rotation of the propeller. The disk brake has an axially extending shaft having at one end a disk and at the opposite end a gear having teeth that engage with one of the gears in the epicyclic gearing arrangement. The disk brake includes a hydraulic caliper or an electric caliper that can be actuated to slow down rotation of the disk to an eventual full stop and to hold the disk at full stop to thereby stop rotation of the propeller.

Growing community of inventors

Marblehead, MA, United States of America

Daniel Waslo