General Electric Company

United States of America As Represented by the Secretary of the Air Force

John Michael Koshoffer

Edward E Ekstedt

Kevin Robert Drake

Narendra Digamber Joshi

Lawrence Butler

Barry Weinstein

Roy E Moyer

Edward I Stamm

James Neil Cooper

Phillip D Napoli

Eric John Kress

David Walter Parry

Adrian M Ablett

Randy Lee Lewis

David Elson Caldwell

Joseph F Savelli

Billy P Samuel

William F Burke

James N Reinhold, Jr

Timothy R Gresla

Richard A Pfefferle

Harold A Larson

Mary C Haas

A gas turbine engine augmentor includes an augmentor outer casingand an annular augmentor linerdisposed radially within and radially separated from the augmentor outer casingto form a generally annular augmentor liner cooling flow pathbetween the augmentor casing and the annular augmentor liner, and an augmentor flow splitter ductdisposed at the upstream end of the augmentor liner and radially between the augmentor outer casingand the annular augmentor linerand defining a splitter flow pathbetween the annular augmentor linerand the annular augmentor flow splitter duct. A method of operating a gas turbofan engine including an augmentor includes directing a portion of fan bypass air into a generally annular augmentor liner cooling flow pathbetween an augmentor outer casingand an annular augmentor liner, directing a portion of the air in the augmentor liner cooling flow pathinto a splitter flow pathdisposed radially outside the augmentor linerand maintaining positive backflow pressure margin between air flow within the splitter flow path and the augmentor liner cooling flow path

Exhaust dust flow splitter system

A turbofan gas turbine engine augmenter includes a fuel/air swirler disposed between an axially extending bypass flowpath and an axially extending exhaust flowpath. A swirler inlet is axially open to and positioned substantially normal to the bypass flowpath and a swirler outlet is open to and positioned substantially parallel to the exhaust flowpath. A swirl chamber within the fuel/air swirler is between the swirler inlet and the swirler outlet. A swirl axis of the fuel/air swirler extends through the swirler outlet substantially normal to the exhaust flowpath. An air swirler may be centered about the swirl axis within the fuel/air swirler. The air swirler may be a louvered or have a plurality of swirling vanes. The swirler inlet may be radially offset with respect to the swirl axis. An air scoop may lead from the swirler inlet to a rounded swirler housing within which the air swirler is disposed.

Augmenter swirler pilot

An afterburner includes an exhaust duct with a fixed area outer nozzle at an aft end thereof, and an ablative inner nozzle therein. The smaller inner nozzle is consumed during reheat operation to expose the larger outer nozzle for increased propulsion thrust.

Ablative afterburner

A trapped vortex cavity afterburner includes one or more trapped vortex cavity stages for injecting a fuel/air mixture into a combustion zone. The trapped vortex cavity afterburner is operable to provide all thrust augmenting fuel used for engine thrust augmentation. Each stage has at least one annular trapped vortex cavity. The trapped vortex cavity afterburner may be a multi-stage afterburner having two or more trapped vortex cavity stages ganged for simultaneous ignition or operable for sequential ignition. One embodiment of the annular trapped vortex cavity is operable to raise a temperature of an exhaust gas flow through the afterburner about 100 to 200 degrees Fahrenheit. Each of the trapped vortex cavity stages may be operable to produce a single or a different amount of temperature rise of the exhaust gas flow through the afterburner. A chevron shaped trapped vortex cavity and having zig-zag shaped leading and trailing edges may be used.

Trapped vortex cavity afterburner

A turbine rear frame includes a row of outlet guide vanes extending between outer and inner bands. Each of the vanes is bifurcated into a forward prow integrally joined to an aft stern by a septum therebetween. The prow and stern collectively define the aerodynamic profile of each vane which is locally interrupted at the septum.

Bifurcated outlet guide vanes

A method for fabricating an augmentor includes fabricating an outer casing having at least one channel defined therein, fabricating a centerbody having at least one channel defined therein, fabricating a plurality of turbine frame vanes, wherein each turbine frame vane includes a first sidewall having a first channel defined therein, a second sidewall having a second channel defined therein, and at least one cross-fire tube extending between the first sidewall and the second sidewall, and coupling the plurality of turbine frame vanes to the augmentor outer casing and to the centerbody such that the first sidewall first channel formed on a first turbine frame vane, the second sidewall second channel formed on a second turbine frame vane, the augmentor channel, and the centerbody channel define a substantially contiguous trapped vortex chamber.

Method and apparatus for gas turbine engines

A method facilitates generating thrust from a gas turbine engine using a pulse detonation system. The method includes introducing fuel and air to the engine, mixing fuel and air in a pulse detonation system deflagration chamber positioned radially outward from an engine exhaust centerbody, and detonating the fuel and air mixture within the pulse detonation system to facilitate increasing the temperature and pressure within the engine and to generate engine thrust.

Methods for operating gas turbine engines

A method enables thrust to be generated from a gas turbine engine using a pulse detonation system is provided. The engine includes an inlet portion and an exhaust portion, and the pulse detonation system includes a multi-staged pulse detonation augmentor including predetonator. The method comprises supplying a less than stoichiometric fuel/air mixture to the pulse detonation system during a first operating stage, detonating the fuel/air mixture with the predetonator to increase the temperature and pressure within the engine and to generate engine thrust, and supplying additional fuel and air to the pulse detonation system during a second operating stage.

Methods and apparatus for generating gas turbine engine thrust with a pulse detonation thrust augmenter

Methods and apparatus for operating gas turbine engines

A vane assembly for a gas turbine engine includes at least one vane that includes a first body, a second body, and a passageway. The first body includes a first sidewall and a second sidewall that are connected at a leading edge and a trailing edge. The passageway extends between the second body and the first body leading edge.

Method and apparatus for reducing distortion losses induced to gas turbine engine airflow

A control system for selectively adjusting effective flow areas of an aircraft engine exhaust nozzle to change operational characteristics of the nozzle. The control system includes a chamber having a hollow interior, a plurality of outlet passages extending from the hollow interior of the chamber to sites within the exhaust nozzle, and an adjustable inlet extending from a pressurized air source to the hollow interior of the chamber for delivering a jet of pressurized air to the chamber. The inlet is adjustable to direct pressurized air to selected one or more outlet passages for the delivery of air via the one or more outlets to corresponding one or more sites within the exhaust nozzle, thereby to change the operational characteristics of the nozzle.

Fluidic nozzle control system

A gas turbine engine of the type generating a hot gas stream and a fan stream includes a mixer having upstream and downstream portions which define first and second chutes for efficiently mixing the streams prior to discharge through an engine nozzle. Means are provided to selectively communicate the downstream chutes with alternate ones of the upstream chutes to cool the downstream mixer portion and reduce the infrared radiation emissions from the nozzle.

Infrared radiation suppression device

An annular combustor with an air fuel mixer is disclosed including a plurality of mixing tubes having a forward end and an exit end, wherein a longitudinal axis through the exit end for each mixing tube is at both an axial angle and a radial angle to the centerline axis of the mixer. A fuel injector for providing fuel to the forward end of each mixing tube, a fuel manifold in flow communication with each of the fuel injectors, and a fuel supply and control means are provided. High pressure air from a compressor is injected into the mixing tubes and fuel is injected into the mixing tubes from the fuel injectors, wherein the high pressure air and the fuel is uniformly mixed therein so as to produce minimal formation of pollutants when the fuel/air mixture is exhausted out the downstream end of the mixing tubes into the combustor and combusted. The mixing tubes preferably impart a swirl to the fuel/air mixture as it exits into the combustor.

Gas turbine combustor with a plurality of circumferentially spaced

A gas turbine engine of the type generating a hot gas stream and a fan stream includes a mixer having upstream and downstream portions which define first and second chutes for efficiently mixing the streams prior to discharge through an engine nozzle. Means are provided to selectively communicate the downstream chutes with alternate ones of the upstream chutes to cool the downstream mixer portion and reduce the infrared radiation emissions from the nozzle,

In accordance with one aspect of the present invention, a combustor is disclosed having three separate domes, each dome having disposed therein a plurality of circumferentially spaced mixers for uniformly mixing air and fuel. The combustor utilizes a high percentage of the total available combustion air at the dome area in order to maintain a lean mixture of air and fuel during all levels of engine output. Besides providing an unobstructed area in the diffuser dump region upstream of the mixers for the combustion air, sealing means is included to prevent cooling air from entering the combustion chamber until downstream of the primary combustion zone. Heat shields having radial centerbodies are also included between each dome to isolate each primary combustion zone.

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Cincinnati, OH, United States of America

John Michael Koshoffer