Siemens Energy, Inc.

Mikro Systems, Inc.

Siemens Aktiengesellschaft

Mikro Sysytems, Inc.

Rtx Corporation

John Joseph Marra

Ching-Pang Lee

Christian X Campbell

Jay Alan Morrison

Gary Brian Merrill

Allister William James

Jan H Marsh

Nan Jiang

Samuel R Miller, Jr

Brian J Wessell

Kok-Mun Tham

Benjamin E Heneveld

Eric Joseph Schroeder

Darryl S Eng

Humberto A Zuniga

Jill Klinger

Ronald J Rudolph

Brede J Kolsrud

Jamie Meeroff

George P Liang

Paul Hadley Vitt

Paul John Zombo

Daniel M Eshak

Clinton A Mayer

Stefan Mazzola

Raymond G Snider

Paul J Gear

John P Dalton

Andrew Whalley

David A Kemp

Andrew S Lohaus

Christian X Campbell, Jr

An attachment system () for a turbine airfoil () including one or more roots () to attach the turbine airfoil () to a rotor, whereby the roots () have one or more curved teeth () that substantially limit, if not completely eliminate, circumferential rocking motion of the turbine airfoil () relative to a disc () supporting the turbine airfoil () during turning gear operation. The curved configuration of the teeth () extending laterally from the root () prevent rotation of the turbine airfoil () relative to the disc () supporting the turbine airfoil (), thereby preventing premature failure of the turbine airfoil () or disc () due to wear from turbine airfoil () rocking during turning gear operation. In at least one embodiment, a laterally extending outer edge () of an axially extending tooth () may be curved about an axis () orthogonal to a centerline () of a turbine engine () in which the turbine airfoil () is positioned.

Attachment system for a turbine airfoil usable in a gas turbine engine

A cooling system for a turbine engine for directing cooling fluids from a compressor to a turbine blade cooling fluid supply and from an ambient air source to the turbine blade cooling fluid supply to supply cooling fluids to one or more airfoils of a rotor assembly is disclosed. The cooling system may include a compressor bleed conduit extending from a compressor to the turbine blade cooling fluid supply that provides cooling fluid to at least one turbine blade. The compressor bleed conduit may include an upstream section and a downstream section whereby the upstream section exhausts compressed bleed air through an outlet into the downstream section through which ambient air passes. The outlet of the upstream section may be generally aligned with a flow of ambient air flowing in the downstream section. As such, the compressed air increases the flow of ambient air to the turbine blade cooling fluid supply.

Cooling system with compressor bleed and ambient air for gas turbine engine

A vane assembly () having: an airfoil () and a shroud () held together without metallurgical bonding there between; a channel () disposed circumferentially about the airfoil (), between the airfoil () and the shroud (); and a seal () disposed in the channel (), wherein during operation of a turbine engine having the vane assembly () the seal () has a sufficient ductility such that a force generated on the seal () resulting from relative movement of the airfoil () and the shroud () is sufficient to plastically deform the seal ().

Ductile alloys for sealing modular component interfaces

A cooling channel (B) cools an exterior surface (or) or two opposed exterior surfaces (and). The channel has a near-wall inner surface () with a width (W1). Interior side surfaces () may converge to a reduced channel width (W2). The near-wall inner surface () may have fins () aligned with a coolant flow (). The fins may highest at mid-width of the near-wall inner surface. A two-sided cooling channel () may have two near-wall inner surfaces () parallel to two respective exterior surfaces (), and may have an hourglass shaped transverse sectional profile. The tapered channel width (W1, W2) and the fin height profile (A,B) increases cooling flow () into the corners (C) of the channel for more uniform and efficient cooling.

Component cooling channel

Platforms () span between turbine blades () on a disk (). Each platform may be individually mounted to the disk by a pin attachment (). Each platform () may have a rotationally rearward edge portion () that underlies a forward portion () of the adjacent platform (). This limits centrifugal bending of the rearward portion of the platform, and provides coolant sealing. The rotationally forward edge (A,B) of the platform overlies a seal element () on the pressure side () of the forwardly adjacent blade, and does not underlie a shelf on that blade. The pin attachment allows radial mounting of each platform onto the disk via tilting () of the platform during mounting to provide mounting clearance for the rotationally rearward edge portion (). This facilitates quick platform replacement without blade removal.

Turbine blade and non-integral platform with pin attachment

A vane array adapted to be coupled to a vane carrier within a gas turbine engine is provided comprising: a plurality of elongated airfoils comprising at least a first airfoil and a second airfoil located adjacent to one another; a U-ring; first connector structure for coupling a radially inner end section of each of the first and second airfoils to the U-ring; second connector structure for coupling a radially outer end section of each of the first and second airfoils to the vane carrier; a platform extending between the first and second airfoils; and platform connector structure for coupling the platform to one of the U-ring and the vane carrier.

Vane array with one or more non-integral platforms

A cooling arrangement in a gas turbine system (). The arrangement includes a plurality of flow network units () to transfer heat to cooling fluid, at least one unit including first (), second (), and third () flow sections between openings () in a first wall () and an opening in a second wall () to pass cooling fluid through the walls. The first section includes first flow paths, between the openings in the first wall and the second section, extending to the second section. The third section includes third flow paths, between the second section and the opening in the second wall, to effect flow of cooling fluid. The second section includes one or more cooling fluid flow paths between the first section and the third section. The number of flow paths in the second section is fewer than the number of first flow paths and fewer than the number of third flow paths.

Cooling module design and method for cooling components of a gas turbine system

A turbine airfoil usable in a turbine engine and having at least one ambient air cooling system is disclosed. At least a portion of the cooling system may include one or more cooling channels configured to receive ambient air at about atmospheric pressure. The ambient air cooling system may have a tip static pressure to ambient pressure ratio of at least 0.5, and in at least one embodiment, may include a tip static pressure to ambient pressure ratio of between about 0.5 and about 3.0. The cooling system may also be configured such that an under root slot chamber in the root is large to minimize supply air velocity. One or more cooling channels of the ambient air cooling system may terminate at an outlet at the tip such that the outlet is aligned with inner surfaces forming the at least one cooling channel in the airfoil to facilitate high mass flow.

Turbine airfoil with ambient cooling system

A damping structure for a turbomachine rotor. The damping structure including an elongated snubber element including a first snubber end rigidly attached to a first blade and extending toward an adjacent second blade, and an opposite second snubber end positioned adjacent to a cooperating surface associated with the second blade. The snubber element has a centerline extending radially inwardly in a direction from the first blade toward the second blade along at least a portion of the snubber element between the first and second snubber ends. Rotational movement of the rotor effects relative movement between the second snubber end and the cooperating surface to position the second snubber end in frictional engagement with the cooperating surface with a predetermined damping force determined by a centrifugal force on the snubber element.

Turbine blade damping device with controlled loading

A continuous serpentine cooling circuit forming a progression of radial passages (A,A) between pressure and suction side walls () in a MID region of a turbine airfoil (). The circuit progresses first axially, then tangentially, ending in a last radial passage (A) adjacent to the suction side () and not adjacent to the pressure side (). The passages of the axial progression () may be adjacent to both the pressure and suction side walls of the airfoil. The next to last radial passage (A) may be adjacent to the pressure side wall and not adjacent to the suction side wall. The last two radial passages (A,A) may be longer along the pressure and suction side walls respectively than they are in a width direction, providing increased direct cooling surface area on the interiors of these hot walls.

Integrated axial and tangential serpentine cooling circuit in a turbine airfoil

An outer rim seal arrangement (), including: an annular rim () centered about a longitudinal axis () of a rotor disc (), extending fore and having a fore-end (), an outward-facing surface (), and an inward-facing surface (); a lower angel wing () extending aft from a base of a turbine blade () and having an aft end () disposed radially inward of the rim inward-facing surface to define a lower angel wing seal gap (); an upper angel wing () extending aft from the turbine blade base and having an aft end () disposed radially outward of the rim outward-facing surface to define a upper angel wing seal gap (); and guide vanes () disposed on the rim inward-facing surface in the lower angel wing seal gap. Pumping fins () may be disposed on the upper angel wing seal aft end in the upper angel wing seal gap.

Aft outer rim seal arrangement

An airfoil in a gas turbine engine includes an outer wall and an inner wall. The outer wall includes a leading edge, a trailing edge opposed from the leading edge in a chordal direction, a pressure side, and a suction side. The inner wall is coupled to the outer wall at a single chordal location and includes portions spaced from the pressure and suction sides of the outer wall so as to form first and second gaps between the inner wall and the respective pressure and suction sides. The inner wall defines a chamber therein and includes openings that provide fluid communication between the respective gaps and the chamber. The gaps receive cooling fluid that provides cooling to the outer wall as it flows through the gaps. The cooling fluid, after traversing at least substantial portions of the gaps, passes into the chamber through the openings in the inner wall.

Cooled airfoil in a turbine engine

A turbine airfoil usable in a turbine engine and formed from at least an outboard section and an inboard section such that an inner end of the outboard section is attached to an outer end of the inboard section. The outboard section may be configured to provide a tip having adequate thickness and may extend radially inward from the tip with a generally constant cross-sectional area. The inboard section may be configured with a tapered cross-sectional area to support the outboard section.

Turbine airfoil having outboard and inboard sections

A gas turbine engine including: an ambient-air cooling circuit () having a cooling channel () disposed in a turbine blade () and in fluid communication with a source () of ambient air: and an pre-swirler (), the pre-swirler having: an inner shroud (); an outer shroud (); and a plurality of guide vanes (), each spanning from the inner shroud to the outer shroud. Circumferentially adjacent guide vanes () define respective nozzles () there between. Forces created by a rotation of the turbine blade motivate ambient air through the cooling circuit. The pre-swirler is configured to impart swirl to ambient air drawn through the nozzles and to direct the swirled ambient air toward a base of the turbine blade. The end walls () of the pre-swirler may be contoured.

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Winter Springs, FL, United States of America

John Joseph Marra