Power Systems Mfg., LLC

Alstom Technology Limited

Ansaldo Energia IP UK Limited

Ansaldo Energia Switzerland Ag

Gregory Edwin Vogel

Joshua R McNally

Jeremy Metternich

Edwin John Kawecki

Russell B Jones

Hartmut Haehnle

Remigi Tschuor

Alex Torkaman

Elena P Pizano

Vladimir Kitaigorod

Hany Rizkalla

Shailendra Naik

Douglas James Dietrich

Gene Chong

Chad Garner

Matthew Yaquinto

Stephen Wayne Fiebiger

Fred Hernandez

Wes Smith

Ramesh Keshava-Bhattu

Richard Whiting

Bryan Kalb

David E Paredes

Christopher Johnston

Jacob Snyder

Michele Borja

Thomas Rosenbarger

Jeff Tessier

Jeffrey E Tessier

Alain Hernandez

Marion Duggans (-Oneil)

James Brackett

Bernard Tam Yen Sam

An injector for a combustor of a gas turbine engine is provided with a plurality of air/fuel mixing tubes divided into radially outer and radially inner subsets of air/fuel mixing tubes with a first fuel manifold in fluid communication with the radially outer subset of air/fuel mixing tubes and a second fuel manifold in fluid communication with the radially inner subset of air/fuel mixing tubes. Each of the air/fuel mixing tubes of the radially outer subset of air/fuel mixing tubes includes a substantially quadrilateral cross-sectional profile, and each of the air/fuel mixing tubes of the radially inner subset of air/fuel mixing tubes includes a substantially circular cross-sectional profile.

Multitube pilot injector having a split airflow for a gas turbine engine

Modular assemblies for gas turbine engines such as modular vane assemblies and methods of manufacturing the same. The modular assembly includes a first modular component such as a vane platform having a first mating pocket, and a second modular such as an airfoil. The second modular component includes circumferentially extending first and second surfaces at first and second distal ends thereof, respectively, with the first surface being received within the first pocket when the modular assembly is in the assembled state. The second modular component also includes a coating pocket extending from the first surface to the second surface. The coating pocket is recessed towards an interior of the second modular component with respect to first surface and the second surface, and a thermal barrier coating is included within the coating pocket and not included on the first surface or the surface.

Modular components for gas turbine engines and methods of manufacturing the same

An airfoil and turbine vanes and vane assemblies incorporating the same. The airfoil has an uncoated profile substantially in accordance with Cartesian coordinate values of X, Y, and Z as set forth in Table 1, carried to four decimal places. The Z values refer to a percentage of the radial span of the airfoil measured radially from a radially outwardly facing surface of the inner platform. The turbine vane includes an inner platform, an outer platform, and an airfoil such as the one discussed above extending radially outward from the inner platform toward the outer platform. And the vane assembly includes an inner platform, an outer platform, and two or more first stage vanes extending from the inner platform to the outer platform. Each of the two or more first stage vanes include an airfoil as discussed above.

Additive manufacturing optimized first stage vane

Airfoils, gas turbine engine assemblies including such airfoils, and methods of manufacturing the same. The airfoil includes multiple cooling channels proximate a leading edge of the airfoil, each of the cooling channels including an inlet provided on one an inner surface of the airfoil in one of the pressure side wall and the suction side wall, and an outlet provided on an outer surface of the airfoil in the other of the pressure side wall and the suction side wall. The cooling channels are arranged in a staggered configuration such there is an alternate pattern of cooling fluid flow provided at the leading edge of the airfoil, near the airfoil's stagnation point.

Near wall leading edge cooling channel for airfoil

A method for manufacturing and heat treating a component formed from an alloy or superalloy. The component may be part of a gas turbine engine, such as an airfoil of a gas turbine engine turbine, or a combustor fuel nozzle, that includes internal cooling passages. At least part of the component is formed from the alloy or superalloy using an additive manufacturing process. The method includes determining an incipient melting point range and a recrystallization temperature of the alloy or superalloy. The component is then heated to at least the recrystallization temperature and within the incipient melting point range for a predetermined time, and ultimately cooled.

Heat treatment process for additive manufactured components

The present invention includes systems and methods for providing cooling channels located within walls of a turbine airfoil. These cooling channels include micro-circuits that taper in various directions along the length and width of the airfoil. In addition, these cooling channels have a variety of shapes and areas to facilitate convective heat transfer between the surrounding air and the airfoil.

Tapered cooling channel for airfoil

Methods and systems of cooling airfoils are provided. The present invention provides systems and methods for providing cooling channels located within walls of a turbine airfoil. These cooling channels include micro-circuits that taper in various directions along the length and width of the airfoil. In addition, these cooling channels have a variety of shapes and areas to facilitate convective heat transfer between the surrounding air and the airfoil.

Leading edge cooling channel for airfoil

An airfoil for a gas turbine engine, the airfoil comprising a wall having a first surface, a second surface, and a passageway extending through the wall from a first opening in the first surface to a second opening in the second surface, the passageway having one or more sections between the first opening and the second opening, the one or more sections in fluid communication with each other, the plurality of sections comprising a first diffuser section providing a first change in cross-sectional area within the passageway, a second diffuser section providing a second change in cross-sectional area within the passageway, a flow conditioning section, and an edge section having two surfaces set opposite each other across the passageway, the two surfaces extending along the passageway substantially in parallel to one another, the edge section being located adjacent to the second opening.

Airfoil cooling passageways for generating improved protective film

An apparatus for providing improved cooling to a combustion liner of a gas turbine combustor is provided. A plurality of flow deflectors is secured to a flow sleeve in order to improve the flow of impingement air from a flow sleeve to the combustion liner outer surface, such that the amount of impingement air being swept away by a cross flow of cooling air is reduced.

Flow sleeve deflector for use in gas turbine combustor

This disclosure describes a removable flow control device which may be used in a rotating flow supply system in a gas turbine to optimize coolant flow by improving flow dynamics, reducing leakage of coolant, and reducing pressure loss in the flow supply system. The flow control device may be coupled to a blade and rotor assembly and may include a flow modifier for directing flow through a junction at which cooling channels intersect and are in fluid communication. The device may direct, control, meter, channel, or otherwise modify the flow of coolant, and may be coupled to the blade and rotor assembly independently of other blade components so that coupling and decoupling the flow control device does not require modification or de-stacking of the rotor assembly.

Flow control device for rotating flow supply system

A system and method of extending the useable life of a gas turbine blade is disclosed in which the gas turbine blade includes an undercut configuration designed to relieve mechanical and thermal stress imparted into the pedestal region of the airfoil trailing edge. The embodiments of the present invention include turbine blade configurations having different trailing edge undercut configurations as well as additional cooling supplied to the internal passages of the trailing edge region of the turbine blade.

Turbine blade platform undercut

A thermal machine includes a hot gas channel; a shell bounding the hot gas channel; a cooling shirt surrounding the shell; and a cooling channel disposed between the shell and the cooling shirt and configured to convection cool the hot gas channel with a cooling medium, wherein the cooling shirt includes at least one local divergence in the guidance of the cooling medium so as to compensate for non-uniformities in at least one of a thermal load on the shell and a flow of the cooling medium in the cooling channel.

Gas turbine having an improved cooling architecture

A thermal machine including a wall defining a hot gas duct for transferring a hot gas stream and a cooling jacket disposed at a distance from the wall on an outside of the hot gas duct so as to define a cooling duct with an inlet and an outlet. The cooling duct is configured to conduct a cooling medium along an external face of the wall from the inlet to an outlet in a direction counter to a flow of hot gas in the hot gas duct. An impingement cooling plate is disposed at the inlet of the cooling duct and includes cooling baffle holes configured such that cooling medium entering the cooling duct through the cooling baffle holes flows in a direction perpendicular to the wall. The impingement cooling plate is positioned such that an inflow-side edge sealingly abuts the wall of the hot gas duct so as to reduce a transverse flow of the cooling medium through the cooling duct.

Impingement cooling plate for a hot gas duct of a thermal machine

A gas turbine airfoil () includes a pressure sidewall () and a suction sidewall (), extending from a root to a tip and from a leading edge region to a trailing edge and having at least one cooling passage between the pressure sidewall () and the suction sidewall () for cooling air to pass through and cool the airfoil from within. One or several of the cooling passages () extend along the leading edge of the airfoil () and several film cooling holes () extend from the internal cooling passages () along the leading edge region to the outer surface of the leading edge region. The film cooling holes () each have a shape that is diffused in a radial outward direction of the leading edge of the airfoil () at least over a part of the length of the film cooling hole (). Improved cooling in the leading edge region can be achieved because the cooling holes () have a principal axis (), and the shape is asymmetrically diffused in that it is diffused in the radial outward direction from the principal axis () along a forward inclination axis (), and it is additionally diffused in a second lateral direction from the principal axis () along a lateral inclination axis ().

Growing community of inventors

Palm Beach Gardens, FL, United States of America

Gregory Edwin Vogel