Siemens Aktiengesellschaft

Siemens Energy, Inc.

Siemens Energy Global Gmbh & Co. Kg

Jae Y Um

Ching-Pang Lee

Steven Koester

Eric Joseph Schroeder

Mrinal Munshi

Zhengxiang Pu

Anthony Waywood

Gerald L Hillier

Harry Holloman

Gm Salam Azad

Sin Chien Siw

Humberto A Zuniga

Wayne J McDonald

Mohamed Abdullah

Darryl S Eng

Nan Jiang

Christopher K Rawlings

Erik Johnson

Friedrich T Rogers

Thomas F Pechette

Ralph W Matthews

Darryl Eng

Thomas Pechette

A turbine airfoil includes a trailing edge coolant cavity between a pressure sidewall and a suction sidewall. The trailing edge coolant cavity is positioned adjacent to and extending out to a trailing edge of the turbine airfoil. The interior includes an internal arrangement comprising an array of discrete fins formed aft of the trailing edge coolant cavity along the trailing edge.

Turbine airfoil with trailing edge features and casting core

A blade airfoil for a turbine engine that includes an internal multiple pass serpentine flow cooling circuits with a leading edge circuit and a trailing edge circuit. The entrance of a cavity in the leading edge circuit has a narrowing of a cavity width that expands further downstream to a consistent cavity width similar to the cavity width of the rest of the leading edge circuit.

Turbine rotor airfoil and corresponding method for reducing pressure loss in a cavity within a blade

A core structure () includes a first core element () including a leading edge section (), a tip section (), and a turn section () joining the leading edge and tip sections (). The first core element () is adapted to be used to form a leading edge cooling circuit () in a gas turbine engine airfoil (). The leading edge cooling circuit () includes a cooling fluid passage () having a leading edge portion () formed by the first core element leading edge section (), a tip portion () formed by the first core element tip section (), and a turn portion () formed by the first core element turn section (). Each of the leading edge portion (), the tip portion (), and the turn portion () of the cooling fluid passage () are formed concurrently in the airfoil () by the first core element ().

Gas turbine airfoil including integrated leading edge and tip cooling fluid passage and core structure used for forming such an airfoil

A shroud cooling system configured to cool a shroud adjacent to an airfoil within a gas turbine engine is disclosed. The turbine engine shroud may be formed from shroud segments that include a plurality of cooling air supply channels extending through a forward shroud support for impingement of cooling air onto an outer radial surface of the shroud segment with respect to the inner turbine section of the turbine engine. The channels may extend at various angles to increase cooling efficiency. The backside surface may also include various cooling enhancement components configured to assist in directing, dispersing, concentrating, or distributing cooling air impinged thereon from the channels to provide enhanced cooling at the backside surface. The shroud cooling system may be used to slow down the thermal response by isolating a turbine vane carrier from the cooling fluids while still providing efficient cooling to the shroud.

Shroud cooling system for shrouds adjacent to airfoils within gas turbine engines

An airfoil () for a gas turbine engine in which the airfoil () includes an internal cooling system () with one or more internal cavities having an insert () contained within an aft cooling cavity () to form nearwall cooling channels having enhanced flow patterns is disclosed. The flow of cooling fluids in the nearwall cooling channels may be controlled via a plurality of cooling fluid flow controllers () extending from the outer wall () forming the generally hollow elongated airfoil (). The cooling fluid flow controllers () may be collected into spanwise extending rows. In at least one embodiment, the cooling fluid flow controllers () may be positioned within a pressure side nearwall cooling channel () and a suction side nearwall cooling channel () that are both in fluid communication with a trailing edge channel (). The trailing edge channel () may also include cooling fluid flow controllers () extending between the outer walls () forming the pressure and suction sides.

Internal cooling system with insert forming nearwall cooling channels in an aft cooling cavity of an airfoil usable in a gas turbine engine

An airfoil () for a gas turbine engine in which the airfoil () includes an internal cooling system () with one or more internal cavities () having an insert () contained therein that forms nearwall cooling channels () having enhanced flow patterns is disclosed. The flow of cooling fluids in the nearwall cooling channels () may be controlled via a plurality of cooling fluid flow controllers () extending from the outer wall () forming the generally hollow elongated airfoil (). The cooling fluid flow controllers () may be collected into spanwise extending rows (), and the internal cooling system () may include one or more bypass flow reducers () extending from the insert () toward the outer wall () to direct the cooling fluids through the channels () created by the cooling fluid flow controllers (), thereby increasing the effectiveness of the internal cooling system ().

Internal cooling system with insert forming nearwall cooling channels in midchord cooling cavities of a gas turbine airfoil

A turbine rotor blade includes at least two integrated cooling circuits that are formed within the blade that include a leading edge circuit having a first cavity and a second cavity and a trailing edge circuit that includes at least a third cavity located aft of the second cavity. The trailing edge circuit flows aft with at least two substantially 180-degree turns at the tip end and the root end of the blade providing at least a penultimate cavity and a last cavity. The last cavity is located along a trailing edge of the blade. A tip axial cooling channel connects to the first cavity of the leading edge circuit and the penultimate cavity of the trailing edge circuit. At least one crossover hole connects the penultimate cavity to the last cavity substantially near the tip end of the blade.

Integrated circuit cooled turbine blade

A component in a gas turbine engine includes an airfoil extending radially outwardly from a platform associated with the airfoil. The airfoil includes opposed pressure and suction sidewalls, which converge at a first location defined at a leading edge of the airfoil and at a second location defined at a trailing edge of the airfoil opposed from the leading edge. The component includes a built-up surface adjacent to the leading edge at an intersection between the pressure sidewall and the platform, and at least one cooling passage at least partially within the built-up surface at the intersection between the pressure sidewall and the platform. The at least one cooling passage is in fluid communication with a main cooling channel within the airfoil and has an outlet at the platform for providing cooling fluid directly from the main cooling channel to the platform.

Airfoil having built-up surface with embedded cooling passage

A film cooling structure formed in a component wall of a turbine engine and a method of making the film cooling structure. The film cooling structure includes a plurality of individual diffusion sections formed in the wall, each diffusions section including a single cooling passage for directing cooling air toward a protuberance of a wall defining the diffusion section. The film cooling structure may be formed with a masking template including apertures defining shapes of a plurality of to-be-formed diffusion sections in the wall. A masking material can be applied to the wall into the apertures in the masking template so as to block outlets of cooling passages exposed through the apertures. The masking template can be removed and a material may be applied on the outer surface of the wall such that the material defines the diffusion sections once the masking material is removed.

Component wall having diffusion sections for cooling in a turbine engine

An investment casting method for a cast ceramic core (), including an airfoil portion () shaped to define an inner surface () of an airfoil () of a vane segment () and an integral shell portion () having a backside-shaping surface () shaped to define a backside surface () of a shroud () of the vane segment. The backside-shaping surface has a higher elevation () and a lower elevation (). The higher elevation is set apart from a nearest point () on the airfoil portion by the lower elevation. The airfoil portion and the shell portion are cast as a monolithic body during a single casting pour.

Investment casting method for gas turbine engine vane segment

A component wall in a turbine engine. The component wall includes a substrate, a trench, and a plurality of cooling passages. The substrate has a first surface and a second surface opposed from the first surface. The trench is located in the second surface and is defined by a bottom surface between the first and second surfaces, a first sidewall, and a second sidewall spaced from the first sidewall. The first sidewall extends radially outwardly continuously from the bottom surface of the trench to the second surface. The first sidewall includes a plurality of first protuberances extending toward the second sidewall. The cooling passages extend through the substrate from the first surface to the bottom surface of the trench. Outlets of the cooling passages are arranged within the trench such that cooling air exiting the cooling passages is directed toward respective ones of the first protuberances.

Growing community of inventors

Winter Garden, FL, United States of America

Jae Y Um