Abstract: A blade includes an airfoil defined by a pressure side outer wall and a suction side outer wall connecting along leading and trailing edges and form a radially extending chamber for receiving a coolant flow. A rib configuration may include: a leading edge transverse rib connecting the pressure side outer wall and the suction side outer wall and partitioning the radially extending chamber into a leading edge passage within the leading edge of the airfoil and a central passage adjacent to the leading edge passage. One or both camber line ribs connect to a corresponding pressure side outer wall and suction side outer wall at a point aft of the leading edge transverse rib causing the central passage to extend towards one or both of the pressure side outer wall and the suction side outer wall, resulting in a flared center cavity aft of the leading edge.

Abstract: A turbine blade includes an airfoil defined by a pressure side outer wall and a suction side outer wall connecting along leading and trailing edges and form a radially extending chamber for receiving a coolant flow. A rib configuration may include: a leading edge transverse rib connecting to the pressure side outer wall and the suction side outer wall and partitioning a leading edge passage from the radially extending chamber. The rib configuration may also include a first center transverse rib connecting to the pressure side outer wall and the suction side outer wall and partitioning an intermediate passage from the radially extending chamber directly aft of the leading edge passage. The intermediate passage is defined by the pressure side outer wall, the suction side outer wall, the leading edge transverse rib and the first center transverse rib, and thus spans airfoil between its outer walls.

Abstract: A blade includes an airfoil defined by a pressure side outer wall and a suction side outer wall connecting along leading and trailing edges and forming a radially extending coolant receiving chamber. A rib partitions the radially extending chamber into a first passage on a first side of the rib and an adjacent second passage on an opposing second side of the rib. Each passage is enclosed at an end of the radially extending chamber by an end member of the radially extending chamber. A turn opening is defined in an end of the rib through which the coolant passes between the first passage and the second passage within the end member of the radially extending chamber. A bulbous projection extends along the end of the rib and on opposing radially extending sides of the turn opening to reduce stress in the rib and/or connecting fillets.

Abstract: An internal rib for a blade airfoil has a concave surface defined to ensure durability and provide desired heat transfer. A concave surface faces a pressure side or suction side outer wall. A width is between a first end and a second end, and a depth is a length of a normal depth line between a midpoint of the concave surface and an intersection point of the depth line with the pressure or suction side outer wall. An irregular arc is defined within an arc angle centered at the intersection point, the irregular arc has a first arc radius equivalent to the depth at the midpoint of the concave surface and a second arc radius where the arc angle intersects the concave surface equivalent to a product of the depth and a shape factor. The shape factor has a substantially linear relationship with the aspect ratio.

Abstract: One aspect of the disclosure provides for a turbine airfoil. The turbine airfoil may include a trailing edge having: a set of cooling channels having a first cooling channel fluidly connected to a second cooling channel; a first section having a first pin bank cooling arrangement, the first section fluidly connected to the first cooling channel; and a second section having a second pin bank cooling arrangement, the second section fluidly connected to the second cooling channel and being radially inward of the first section.

Abstract: A cooling system for a turbine bucket including a multi-wall blade and a platform. A cooling circuit for the multi-wall blade includes: an outer cavity circuit and a central cavity for collecting cooling air from the outer cavity circuit; a platform core air feed for receiving the cooling air from the central cavity; and an air passage for fluidly connecting the platform core air feed to a platform core of the platform.

Abstract: A core for a turbine airfoil casting according to an embodiment includes: a center plenum section; and a plurality of outer passage sections; wherein the center plenum section includes at least one boss extending outwardly from the center plenum to an outer profile of the core.

Abstract: One aspect of the disclosure provides for a turbine airfoil. The turbine airfoil may include a trailing edge having: a set of cooling channels having a first cooling channel fluidly connected to a second cooling channel; a first section having a first pin bank cooling arrangement, the first section fluidly connected to the first cooling channel; a second section having a second pin bank cooling arrangement, the second section fluidly connected to the second cooling channel and being radially inward of the first section; and a pressure side panel having a third pin bank cooling arrangement, the pressure side panel fluidly connected to the first cooling channel.

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: 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.