Siemens Energy, Inc.

Mikro Systems, Inc.

Siemens Aktiengesellschaft

Humberto A Zuniga

Ching-Pang Lee

Jay Alan Morrison

John Joseph Marra

Mrinal Munshi

Jae Y Um

Brede J Kolsrud

Michael E Crawford

An internal cooling system () including an impingement jet strike channel system () for increasing the effectiveness of impingement jets () is disclosed. The impingement jet strike channel system () may include an impingement jet strike cavity () offset from one or more impingement orifices (). A plurality of impingement jet strike channels () may extend radially outward from the impingement jet strike cavity () forming a starburst pattern of impingement jet strike channels () and may be formed by a plurality of ribs () that each separate adjacent impingement jet strike channels (). The ribs () forming the impingement jet strike channels () may be split one or more times into multiple channels to increase the number of stagnation points () to increase the cooling capacity. The impingement jet strike channel system () may be used within components, such as, but not limited to, gas turbine engines (), including vane inserts, airfoil leading edge cooling systems, platforms, advanced transitions, acoustic resonators, ring segments and the like.

Impingement jet strike channel system within internal cooling systems

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 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

A component wall in a turbine engine includes a substrate, a diffusion section, and at least one cooling passage. The diffusion section is located in a surface of the substrate and is defined by a first sidewall and a second sidewall. The cooling passage(s) include an outlet portion through which cooling air exits in a direction toward the first sidewall. The outlet portion includes a rear section, a front section, and an inner wall having proximal and distal ends. The rear section is located between the first and second sidewalls. The front section extends between the first sidewall and the distal end of the inner wall. The first sidewall extends into the outlet portion of the cooling passage(s) to the inner wall and extends from the first lateral wall to the second lateral wall so as to block the front section of the outlet portion.

Cooled component wall in a turbine engine

A conduit through which hot combustion gases pass in a gas turbine engine. The conduit includes a wall structure having an inner surface, an outer surface, a region, an inlet, and an outlet. The inner surface defines an inner volume of the conduit. The region extends between the inner and outer surfaces and includes cooling fluid structure defining a plurality of cooling passageways. The inlet extends inwardly from the outer surface and provides fluid communication between the inlet and the passageways. The outlet extends from the passageways to the inner surface to provide fluid communication between the passageways and the inner volume. At least one first cooling passageway intersects with at least one second cooling passageway such that cooling fluid flowing through the first cooling passageway interacts with cooling fluid flowing through the second cooling passageway.

Mesh cooled conduit for conveying combustion gases

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.

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Casselberry, FL, United States of America

Humberto A Zuniga