Solar Turbines Inc.

Sofar Turbines Incorporated

Juan Yin

Luzeng Zhang

Hee Koo Moon

Hee-Koo Moon

Steven D Gray

Scott L Stafford

Xubin Gu

Gas turbine engines generally comprise a first-stage nozzle guide vane. Temperatures in a trailing-edge area of the suction-side wall of such vanes can exceed material and coating limits. While an insert can be used to form passages for cooling air to flow along the inner surfaces of the vane walls, design constraints prevent the insert from extending beyond a certain point into the trailing edge of the vane. Accordingly, a fin is disclosed for insertion downstream of the insert. By eliminating sudden expansion beyond the downstream end of the insert and maintaining the speed of the cooling air across the trailing-edge area of the suction-side wall, the fin improves the cooling coefficient for the trailing-edge area, so as to prevent or reduce excessive temperatures in the trailing-edge area.

Fin for internal cooling of vane wall

An airfoil for a turbomachine such as a gas turbine engine includes a pressure side and a suction side, a flow-distributing forward wall and an inner cooling wall. A switchbacked passage extends through the airfoil, and the flow-distributing forward wall has a plurality of jetting orifices formed therein and connecting the switchbacked passage to a forward cavity. The jetting orifices are oriented to produce wall jets of cooling fluid directed in an upstream direction toward a back side of the leading edge. A second passage extends between the forward cavity and outlets of the airfoil, so as to convey cooling fluid in a downstream direction toward the outlets.

Airfoil for turbomachine and airfoil cooling method

An airfoil for a turbomachine such as a gas turbine engine includes a switchbacked passage for conveying cooling fluid, and flow-distribution passages in a forward wall so as to direct cooling fluid from the internal passage for back side impingement upon a leading edge of the airfoil. An inner cooling wall forms a passage for discharging the spent cooling air from a trailing edge of the airfoil after impingement, and cools one of a pressure side and a suction side of the airfoil by way of conduction. The switchbacked passage may have a serpentine form.

Airfoil for turbomachine and method of cooling same

A turbine blade for a gas turbine engine may include an outer, peripheral wall extending along a suction side and a pressure side of the blade from a leading edge to a trailing edge of the blade and from a root end to a tip end of the blade, a substantially vertical bifurcating internal wall extending between the leading edge and the trailing edge of the blade in between the suction side and the pressure side of the blade, a plurality of pressure side cooling fluid passages defined between the peripheral wall on the pressure side of the blade and the bifurcating internal wall and extending at least part way from the root end to the tip end of the blade, and a plurality of suction side cooling fluid passages defined between the peripheral wall on the suction side of the blade and the bifurcating internal wall and extending at least part way from the tip end to the root end of the blade. At least one of the pressure side cooling fluid passages may receive fresh cooling fluid from a cooling fluid inlet opening at the root end of the blade and direct the cooling fluid radially upwardly toward the tip end of the blade, across a top side of the bifurcating internal wall, and down through at least one of the plurality of suction side cooling fluid passages.

Rotation enhanced turbine blade cooling

A turbine blade for a gas turbine engine may include at least two wrapped, serpentine-shaped internal cooling paths. A first one of the serpentine-shaped internal cooling paths may include a first passage that extends radially along a leading edge of the turbine blade from adjacent a root end of the turbine blade to adjacent a tip end of the turbine blade. The first passage may be configured to provide fresh cooling fluid to the leading edge. A second passage downstream of the first passage may be configured to discharge spent cooling fluid from the first passage of the first one of the serpentine-shaped internal cooling paths across a plurality of flow disrupters positioned along an upper span of a trailing edge of the turbine blade before exiting from the trailing edge of the turbine blade. A second one of the serpentine-shaped internal cooling paths may be configured to supply fresh cooling fluid to a lower span of the trailing edge of the turbine blade.

Wrapped serpentine passages for turbine blade cooling

A nozzle segment for a nozzle ring of a gas turbine engine is disclosed. The nozzle segment includes an upper endwall, a lower endwall, and an airfoil extending between the upper endwall and the lower endwall. The airfoil includes a pressure side wall, a plurality of inner cooling apertures, and a plurality of outer cooling apertures. The plurality of inner cooling apertures extends through a pressure side wall and is arranged in a first row adjacent the lower endwall. The plurality of outer cooling apertures extends through the pressure side wall and is arranged in a second row adjacent the upper endwall.

Nozzle endwall film cooling with airfoil cooling holes

A turbine blade for a gas turbine engine. The turbine blade includes a base having a blade root, a platform, a cooling air inlet, and a base air passageway. The turbine blade also includes an airfoil section adjoined to the base and having an outer wall, an airfoil air passageway, a plurality of trailing edge slots in fluid communication with the airfoil air passageway and a plurality of directional film holes through the outer wall in fluid communication with the airfoil air passageway. The plurality of directional film holes includes a first portion configured to discharge the cooling air toward a tip end, and a second portion configured to discharge the cooling air toward the platform.

Cooled turbine blade with double compound angled holes and slots

A nozzle arrangement for a gas turbine engine comprising a first housing member and a second housing member. The nozzle arrangement may further include a first nozzle and a second nozzle. Each of the first nozzle and second nozzle may extend between the first housing member and the second housing member so as to form a doublet. A plurality of cooling apertures may be arranged on at least one of the first nozzle, the second nozzle, the first housing member, or the second housing member so as to provide a different degree of first order cooling across the doublet.

Turbine nozzle

A turbine blade for a gas turbine engine is disclosed. The turbine blade can include at least one internal cooling path, and an internal vane disposed in the at least one internal cooling path. The internal vane can include a central portion, a first leg extending in a first direction from the central portion, and a second leg extending in a second direction from the central portion. The central portion can have a thickness greater than a thickness of the first leg or a thickness of the second leg.

Growing community of inventors

San Diego, CA, United States of America

Juan Yin