Abstract: A cooling system for a multi-wall blade according to an embodiment includes: a primary cooling air feed for providing cooling air; and a feed splitter coupled to the primary cooling air feed for splitting the cooling air provided by the primary cooling air feed between a pressure side cooling circuit and a suction side cooling circuit.

Abstract: A core for forming micro channels within a turbine component is provided. The core includes a base comprising a first side and a second side; and a core assembly coupled to the second side. The core assembly further includes a plurality of channel members, wherein each channel member has a first end, a second end, and a channel body coupled to and extending between said first end and said second end. The channel body includes a channel shape configured to form the micro channels within the turbine component.

Abstract: A cooling structure for a stationary blade is provided. The cooling structure may include a first chamber in an endwall of the stationary blade directing a first cooling fluid from the stationary blade to a first cooling circuit, and a second chamber in the endwall of the stationary blade directing a second cooling fluid from the stationary blade to a second cooling circuit different than the first cooling circuit. The first cooling fluid has a lower temperature than the second cooling fluid.

Abstract: The present application and the resultant patent provide a turbine nozzle for a gas turbine engine. The turbine nozzle may include a first nozzle vane, a second nozzle vane, and a platform connecting the first nozzle vane and the second nozzle vane. The platform may include a first cooling passage and a separate second cooling passage defined therein. The first cooling passage may be configured to direct a first flow of cooling fluid in a first direction, and the second cooling passage may be configured to direct a second flow of cooling fluid in a second direction substantially opposite the first direction. The present application and the resultant patent further provide a method for cooling a turbine nozzle of a gas turbine engine.

Abstract: Embodiments of the present disclosure provide a cooling structure for a stationary blade, which can include: an endwall coupled to a radial end of an airfoil, relative to a rotor axis of a turbomachine; and a substantially crescent-shaped chamber positioned within the endwall and radially displaced from a trailing edge of the airfoil, the substantially crescent-shaped chamber receiving a cooling fluid from a cooling circuit, wherein the substantially crescent-shaped chamber extends from a fore section positioned proximal to one of a pressure side surface and a suction side surface of the airfoil to an aft section positioned proximal to the trailing edge of the airfoil and the other of the pressure side surface and the suction side surface of the airfoil, wherein the aft section of the substantially crescent-shaped chamber is in fluid communication with the fore section of the substantially crescent-shaped chamber.

Abstract: In one example, an arcuate segment for a ring-shaped, rotary machine component such as a stator nozzle or bucket shroud, includes a segment body having an end face formed with a circumferentially-facing seal slot adapted to receive a seal extending between the segment body and a corresponding seal slot in an adjacent segment body to seal a radially-extending gap between the adjacent segment bodies. A cooling channel is provided in the segment body in proximity to the seal slot, and is adapted to be supplied with cooling air. A passage extends from the cooling channel into the seal slot, at a location where the cooling air can be supplied to the higher pressure area on the radially-outer side of the seal.

Abstract: A gas turbine stator component includes a composite, segmented ring made up of an annular array of arcuate segments, each having end faces formed with respective seal slots, with radial gaps formed between opposed end faces of adjacent arcuate segments. A seal is located between each pair of opposed seal slots to thereby seal the gaps, and a channel is provided in each of said arcuate segments adapted to be supplied with cooling air, the channel connecting to a passage extending between the channel and a respective one of the seal slots or radial gaps, on a lower-pressure side of the seal.

Abstract: Sealing device for providing seals between adjacent components, and turbomachines utilizing such sealing devices, are provided. A sealing device includes a seal plate insertable between the adjacent components, the seal plate comprising a first face and an opposing second face. The sealing device further includes a plurality of pins extending from one of the first face or the second face, the plurality of pins configured to space the one of the first face or the second face from contact surfaces of the adjacent components.

Abstract: A seal assembly for a rotary machine is provided. The seal assembly includes a shim seal including multiple seal plates forming a box shaped shim seal. The box shaped seal includes a plurality of cuts at two opposing sides or corners for allowing high pressure fluid to occupy the cavity of the box-shaped shim seal. The seal may be inserted within one or more slots between adjacent stator components of the rotary machine.

Abstract: A sealing member, a component including a sealing member, and a method of sealing a hole are disclosed. In an embodiment, the sealing member includes a plug member for occluding a hole in a wall of a passageway. The plug member includes at least one cooling feature disposed on a distal end of the plug member exposed to the passageway.

Abstract: The present application provides seal assemblies having improved flexibility for reducing leakages between adjacent misaligned components of turbomachinery. The seal assemblies include a first outer shim formed of a flexible permeable material and a second outer shim formed of a substantially impervious material. At least the second outer shim is configured for sealing engagement with seal slots of the adjacent components. The seal assemblies may also include at least one of an inner shim and a filler layer positioned between the first and second outer shims. The seal assemblies may be sufficiently flexible to account for misalignment between the adjacent components, sufficiently stiff to meet assembly requirements, and sufficiently robust to meet operating requirements associated with turbomachinery. A turbomachine including the seal assembly is provided.

Abstract: A cooling structure for a stationary blade is provided. The cooling structure may include a first chamber in an endwall of the stationary blade directing a first cooling fluid from the stationary blade to a first cooling circuit, and a second chamber in the endwall of the stationary blade directing a second cooling fluid from the stationary blade to a second cooling circuit different than the first cooling circuit. The first cooling fluid has a lower temperature than the second cooling fluid.

Abstract: The present application provides seal assemblies for reducing leakages between adjacent components of turbomachinery. The seal assemblies may include outer shims, and at least a portion of the outer shims may be substantially impervious. At least one of the outer shims may be configured for sealing engagement with seal slots of the adjacent components. The seal assemblies may also include at least one of an inner shim and a filler layer positioned between the outer shims. The at least one inner shim may be substantially solid and the at least one filler layer may be relatively porous. The seal assemblies may be sufficiently flexible to account for misalignment between the adjacent components, sufficiently stiff to meet assembly requirements, and sufficiently robust to operating meet requirements associated with turbomachinery.

Abstract: A blade cooling circuit feed and exhaust duct and related cooling structure are provided. The feed duct may include a feed chamber having a feed entrance fluidly coupled to a cooling fluid source and a feed exit to an elongate entrance to the cooling circuit, the feed exit including a ramped wall maintaining a flow velocity of the cooling fluid along the elongated entrance to the cooling circuit. The exhaust duct may include a substantially concave exhaust chamber including an exhaust entrance at a wider end of the exhaust chamber and in fluid communication with an elongated exit from the cooling circuit, and an exhaust exit at a narrower end of the exhaust chamber, the exhaust exit including an opening to an exhaust passageway from the exhaust chamber.

Abstract: The present embodiments are generally directed toward systems and methods for cooling one or more shroud segments of a gas turbine engine. For example, in a first embodiment, a shroud segment is provided that is configured to at least partially surround a turbine blade of a turbine engine. The shroud segment includes a body and a microchannel disposed in the body. The microchannel is configured to flow a cooling fluid through the body.

Abstract: A turbine shroud cooling assembly for a gas turbine system includes an outer shroud component disposed within a turbine section of the gas turbine system and proximate a turbine section casing, wherein the outer shroud component includes at least one airway for ingesting an airstream. Also included is an inner shroud component disposed radially inward of, and fixedly connected to, the outer shroud component, wherein the inner shroud component includes a plurality of microchannels extending in at least one of a circumferential direction and an axial direction for cooling the inner shroud component with the airstream from the at least one airway.

Abstract: A core for forming micro channels within a turbine component is provided. The core includes a base comprising a first side and a second side; and a core assembly coupled to the second side. The core assembly further includes a plurality of channel members, wherein each channel member has a first end, a second end, and a channel body coupled to and extending between said first end and said second end. The channel body includes a channel shape configured to form the micro channels within the turbine component.

Abstract: Sealing device for providing seals between adjacent components, and turbomachines utilizing such sealing devices, are provided. A sealing device includes a seal plate insertable between the adjacent components, the seal plate comprising a first face and an opposing second face. The sealing device further includes a plurality of pins extending from one of the first face or the second face, the plurality of pins configured to space the one of the first face or the second face from contact surfaces of the adjacent components.

Abstract: The present application and the resultant patent provide a turbine nozzle for a gas turbine engine. The turbine nozzle may include a first nozzle vane, a second nozzle vane, and a platform connecting the first nozzle vane and the second nozzle vane. The platform may include a first cooling passage and a separate second cooling passage defined therein. The first cooling passage may be configured to direct a first flow of cooling fluid in a first direction, and the second cooling passage may be configured to direct a second flow of cooling fluid in a second direction substantially opposite the first direction. The present application and the resultant patent further provide a method for cooling a turbine nozzle of a gas turbine engine.

Abstract: A gas turbine system is provided that includes a compressor section, a combustor assembly coupled to the compressor section, and a turbine section coupled to the compressor section. At least one of the combustor assembly and the turbine section includes a sealing sub-system for use in sealing between a first component and a second component. A first component defines a first seal member receiving region oriented between a higher-temperature gas region and a cooler-temperature gas region. A second component adjacent the first component defines a second seal member receiving region oriented adjacent the first seal member receiving region. The sealing system includes first and second end walls defined in at least one of the first and second seal member receiving regions. A seal member is oriented within the first and second seal member receiving regions, and includes at least a first layer defining at least a first resilient seal end portion that engages one of the first and second end walls.