General Electric Company

Lawrence Wayne Dunbar

James Edward Johnson

Lawrence Butler

Robert Joseph Orlando

Peter Nicholas Szucs

John Christopher Brauer

Jorge Francisco Seda

Thomas Ory Moniz

John J Ciokajlo

Rollin George Giffin

Rollin G Giffin, Iii

Philip Roger Gliebe

Gregory T Steinmetz

Jeremiah P Wolf

Eric A Ford

Ronald F Petsch

Jorge F Seda

An aircraft gas turbine engine includes high and low pressure turbines having respective counter rotatable low pressure inner and outer rotors with low pressure inner and outer shafts in part rotatably disposed co-axially within a high pressure rotor and drivingly connected to first and second fan blade rows and first and second boosters respectively. A bypass duct radially bounded by a fan casing and an annular radially inner bypass duct wall surrounds the boosters axially located between the first and second fan blade rows. The engine has a high pressure compressor operable to produce an overall pressure ratio in a range of about 40-65 and a fan inlet hub to tip radius ratio in a range between 0.20 and 0.35, a bypass ratio in a range of 5-15, an operational fan pressure ratio in a range of 1.4-2.5, and a sum of operational fan tip speeds in a range of 1000 to 2500 feet per second.

Counter rotating aircraft gas turbine engine with high overall pressure ratio compressor

A pulse detonation system for a turbofan engine including a fan assembly and a turbine sub-system, which includes at least one turbine, is configured to create a temperature rise and a pressure rise within the turbofan engine and to generate thrust for the turbofan engine. The pulse detonation system includes a pulse detonation core assembly comprising at least one detonation chamber configured to detonate a fuel mixture. The pulse detonation core assembly is positioned between the fan assembly and the turbine sub-system.

Combined cycle pulse detonation turbine engine

An aircraft gas turbine engine assembly includes an inter-turbine frame axially located between high and low pressure turbines. Low pressure turbine has counter rotating low pressure inner and outer rotors with low pressure inner and outer shafts which are at least in part rotatably disposed co-axially within a high pressure rotor. Inter-turbine frame includes radially spaced apart radially outer first and inner second structural rings disposed co-axially about a centerline and connected by a plurality of circumferentially spaced apart struts. Forward and aft sump members having forward and aft central bores are fixedly joined to axially spaced apart forward and aft portions of the inter-turbine frame. Low pressure inner and outer rotors are rotatably supported by a second turbine frame bearing mounted in aft central bore of aft sump member. A mount for connecting the engine to an aircraft is located on first structural ring.

Aircraft engine with inter-turbine engine frame supported counter rotating low pressure turbine rotors

A turbofan engine includes a pulse detonation system to create a temperature rise and a pressure rise within the engine to generate thrust from the engine. The system includes a pulse detonation augmentor including a shock tube sub-system. The shock tube sub-system includes a plurality of shock tubes which mix air and fuel introduced to the pulse detonation augmentor and detonate the mixture. The detonation creates hot combustion gases which are directed from the engine to produce thrust for the engine. Alternatively, the system includes a pulse detonation augmentation system that replaces a core engine of a turbo-fan engine.

Combined cycle pulse detonation turbine engine operating method

A split fan for providing lift for an aircraft is described. In one embodiment, a main engine shaft extends to the split fan, and a portion of the split fan is located at an elevation above the main engine shaft and a portion of the split fan is located at an elevation below the main engine shaft. In the one embodiment, the split fan includes an outer casing configured to be mounted to a fuselage of the aircraft, a first stage rotor including a shaft, a gear secured to the shaft, and a plurality of rotor blades radially extending from the shaft. The split fan also includes a second stage rotor including a shaft, a gear secured to the shaft, and a plurality of blades radially extending from the shaft. A drive shaft having a first end and a second end is coupled at the drive shaft first end to the main engine shaft. A drive gear is located at the second end of the drive shaft, and the first stage rotor gear and the second stage rotor gear are coupled to the drive gear so that as the drive gear rotates, the first stage rotor gear rotates in a first direction and the second stage rotor gear rotates in a second direction.

Split rotor shaft driven lift fan

A turbine nozzle for a gas turbine engine subassembly having an axial turbine. The turbine nozzle airfoil-shaped vanes are fixedly attached to one or both of the annular inner and outer casings. Apparatus is provided to vary the vane airflow passage area between the casings and between the pressure side of one vane and the suction side of an adjacent vane. The apparatus varies the area near at least the pressure side of the one vane. Various embodiments of the apparatus include: a small pivotable flap pivotally attached to the outer casing and positioned near the pressure side of the one vane; a small movable flap attached to the outer casing to be movable along the pressure side of the one vane, and a plug member insertable into and withdrawable from the vane airflow passage through the outer casing near at least the pressure side of the one vane.

Jet engine variable area turbine nozzle

The present invention is a unique splitter of a core engine compressor in a double bypass, variable cycle engine. The splitter is extended forward to a second fan section. The extended splitter may divide the rotor blades into outer and inner portions to allow the flow past these portions to act essentially aerodynamically independent of each other in a single or double bypass mode of operation of the variable cycle engine.

Compressor splitter for use with a forward variable area bypass injector

A variable cycle jet engine is provided with a mixing duct assembly which mixes core engine exhaust gas with bypass air when the engine is operating in a turbofan mode and which blocks flow from the core engine and isolates the core engine from the bypass flow when the engine is operating as a ramjet.

Admission mixing duct assembly

An array of chutes is permanently mounted within the flowpath of an exhaust nozzle of an aircraft jet engine for entraining and mixing ambient air with the exhaust gas so as to reduce the noise level of aircraft, particularly during take-off. In order to provide good engine performance during all modes of jet engine operation, a convergent-divergent flap assembly is arranged downstream from the chutes for controlling the nozzle throat and exit areas when the chutes are closed.

Growing community of inventors

Cincinnati, OH, United States of America

Lawrence Wayne Dunbar