Siemens Energy, Inc.

Siemens Energy Global Gmbh & Co. Kg

Other

United Technologies Corporation

Samuel R Miller, Jr

John Joseph Marra

Daniel M Eshak

Daniel Vöhringer

Susanne Kamenzky

Ching-Pang Lee

Kok-Mun Tham

Christian X Campbell

Eric Joseph Schroeder

Jamie Meeroff

Mark F Zelesky

Thomas A Auxier

William L Plank

Nan Jiang

Heinrich Stüer

Stefan Schmitt

Friedrich T Rogers

Yuekun Zhou

Andrew Lohaus

Hubert Paprotna

Andrew S Lohaus

A method is provided for machining an airfoil section () of a turbine blade or vane produced by a casting process. The airfoil section () has an outer wall () delimiting an airfoil interior having one or more internal cooling passages (). The method involves: receiving design data pertaining to the airfoil section (), including a nominal outer airfoil form () and nominal wall thickness (T) data; generating a machining path by determining a target outer airfoil form (), the target outer airfoil form () being generated by adapting the nominal outer airfoil form () such that a nominal wall thickness (T) is maintained at all points on the outer wall around the one or more internal cooling passages () in a subsequently machined airfoil section; and machining an outer surface () of the airfoil section () produced by the casting process according to the generated machining path, to remove excess material to conform to the generated target outer airfoil form ().

Adaptive machining of cooled turbine airfoil

A gas turbine blade having a casted metal airfoil, the airfoil has a main wall defining at least one interior cavity, having a first side wall and a second side wall, which are coupled to each other at a leading edge and a trailing edge, extending in a radial direction from a blade root to a blade tip and defining a radial span from 0% at the blade root to 100% at the blade tip. The main airfoil has a radial span dependent chord length defined by a straight line connecting the leading edge and the trailing edge as well as a radial span dependent solidity ratio of metal area to total cross-sectional area. Solidity ratios in a machined zone of the airfoil from 80% to 85% of span are below 35%, in particular all solidity ratios in the zone.

Gas turbine blade and method for producing such blade

A bladed rotor system includes first and second sets of blades with respective airfoils each having at least one internal cavity. The airfoils of both the first and second sets of blades have identical outer shapes defined by an outer surface of an outer wall of the respective airfoils. The airfoils of the first set of blades are distinguished from the airfoils of the second set of blades by a geometry and/or position of at the least one internal cavity, which is unique to blades of a given set. The natural frequency of a blade of the first set differs from the natural frequency of a blade of the second set by a predetermined amount. The blades of the first set and the second set are alternately arranged in a periodic fashion in said circumferential row, to provide a frequency mistuning to stabilize flutter of the blades.

Mistuning of turbine blades with one or more internal cavities

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

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

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.

Ambient air cooling arrangement having a pre-swirler for gas turbine engine blade cooling

A ring segment for a gas turbine engine includes an outer panel defining a structural body for the ring segment. An outer side of an inner panel is attached to an inner side of the outer panel at an interface, and an inner side of the inner panel defines a portion of a hot gas path through the gas turbine engine. An outer side of the outer panel, opposite from the interface, is in communication with a source of cooling air. A plurality of impingement holes extend through the outer panel from the outer side to the inner side of the outer panel for directing impingement air to the outer side of the inner panel. The outer and inner panels define a plurality of flow channels at the interface for effecting convective cooling of the outer panel along the flow channels between the outer and inner panels.

Ring segment with impingement and convective cooling

An air cooled gas engine turbine blade that includes a plurality of longitudinally spaced cavities adjacent the leading edge of the blade is designed to include angularly disposed impingement passages flowing cooling air into each of the cavities in a direction extending from the root to the tip of the blade and including an annular projection upstream of the impingement passage but adjacent thereto for directing air into the respective cavities with total instead of static pressure. The impingement holes are oriented to align with the film cool holes in the blade surface at the leading edge. Ribs formed between cavities are also oriented to be parallel to the impingement holes.

Cooled turbine blade

A spring clip fastened panel for furniture and the like is disclosed comprising a frame forming a surround about a recess for receiving a panel, said frame having inwardly facing edge surfaces bounding the recess and shaped to define a retaining lip to support a panel in the recess. A panel fits into the recess and is supported against the retaining lip. The edge surfaces include a groove in the plane of the panel, and generally U-shaped spring clips having first legs which are inserted in the grooves along approximately diagonal paths inclined to the plane of the grooves and retained in a stressed condition in the grooves by second legs integrally joined to the first legs and having lips at their ends which press against the edge of the panel thereby retaining the panel in the recess.

Growing community of inventors

Port St. Lucie, FL, United States of America

Samuel R Miller, Jr