A core for casting a turbine blade to form at least one cooling passage in a wing portion of the turbine blade, wherein the wing portion includes a leading edge region and a trailing edge region, and has a streamlined cross-section, the core including: at least one of a first core unit having a shape corresponding to a cooling passage located at the leading edge region and a second core unit spaced apart from the first core unit and having a shape corresponding to a cooling passage located at the trailing edge region, wherein each of the first core unit and the second core unit includes: a plurality of extending portions extending in a longitudinal direction and located substantially parallel to one another; at least one curved portion connecting adjacent ends of the plurality of extending portions; and at least one through-portion located between the plurality of extending portions and having an empty space extending in a width direction of the wing portion.

A casting core and/or an airfoil may comprise a tip flag cavity having a forward pedestal and a first spear pedestal disposed aft of the forward pedestal. A trailing edge discharge cavity may be separated from the tip flag cavity and include a first row of pedestals. The first row of pedestals may comprise a first racetrack pedestal. A second row of pedestals may be disposed aft of the first row of pedestals and include a second racetrack pedestal. A third row of pedestals may be disposed aft of the second row of pedestals and include a circular pedestal. A fourth row of pedestals may be disposed aft of the third row of pedestals and include a second spear pedestal.

An engine component assembly includes at least one cavity that is in communication with a source of cooling air. An insert disposed within the cavity includes a plurality of scoops protruding into a flow of cooling air for directing cooling air through the insert and against an inner surface of the cavity.

A moving blade for a wheel of an aircraft turbine engine, including an aerodynamic aerofoil and an outer heel defining the aerofoil. The heel includes a platform and a first lip that projects from the platform. The first lip is inclined upstream and peripherally along an axis of elongation. The heel includes a row of ribs that are arranged at a distance from each other. The row of ribs extends along the axis of elongation and from the platform up to the first lip. The ribs are arranged upstream of the first lip in such a way as to generate turbulence upstream of first lip.

A turbomachine element including an airfoil set including a plurality of airfoils that are offset from one another in a lateral direction, and upstream from at least one end of each airfoil, a group of a plurality of vortex generator devices that are mutually offset both laterally and axially.

A system for supplying fuel includes a fuel manifold, a water manifold, and a fluid junction between the fuel manifold and the water manifold. A turbulator downstream from the fluid junction receives a fluid flow from the fluid junction. A method for supplying fuel includes flowing fuel through a fuel manifold, flowing water through a water manifold, and combining a portion of the water from the water manifold with the fuel from the fuel manifold to create emulsion fuel. The method further includes flowing the emulsion fuel through a turbulator.

A turbine blade has hollowness, and is provided with a back-side wall of which a portion of the inner wall surface is exposed at the rear edge portion, with cooling air flown along the inner wall surface at the exposed region; and a recess provided in the inner wall surface at the exposed region. The contour of the recess (5) viewed from the normal direction of the inner wall surface of the back-side wall is set to a shape that is symmetrical centered on a reference axis (L) that intersects the flow direction of cooling air, and that broadens along the reference axis (L).

An airfoil (10) for a gas turbine engine in which the airfoil (10) includes an internal cooling system (14) with one or more internal cavities having an insert (18) contained within an aft cooling cavity (76) 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 (22) extending from the outer wall (24) forming the generally hollow elongated airfoil (26). The cooling fluid flow controllers (22) may be collected into spanwise extending rows. In at least one embodiment, the cooling fluid flow controllers (22) may be positioned within a pressure side nearwall cooling channel (48) and a suction side nearwall cooling channel (50) that are both in fluid communication with a trailing edge channel (30). The trailing edge channel (30) may also include cooling fluid flow controllers (22) extending between the outer walls (12, 13) forming the pressure and suction sides.

An airfoil (10) for a gas turbine engine in which the airfoil (10) includes an internal cooling system (14) with one or more internal cavities (16) having an insert (18) contained therein that forms nearwall cooling channels (20) having enhanced flow patterns is disclosed. The flow of cooling fluids in the nearwall cooling channels (20) may be controlled via a plurality of cooling fluid flow controllers (22) extending from the outer wall (24) forming the generally hollow elongated airfoil (26). The cooling fluid flow controllers (22) may be collected into spanwise extending rows (28), and the internal cooling system (14) may include one or more bypass flow reducers (30) extending from the insert (18) toward the outer wall (24) to direct the cooling fluids through the channels (20) created by the cooling fluid flow controllers (22), thereby increasing the effectiveness of the internal cooling system (14).

A gas turbomachine includes a compressor portion, a turbine portion operatively connected to the compressor portion, and a combustor assembly fluidically connected to each of the compressor portion and the turbine portion. A turbine casing includes a body having an outer surface and an inner surface. A clearance control system includes a plurality of fluidically connected fluid channels extending through the turbine casing. The plurality of fluidically connected fluid channels includes a first fluid channel configured to direct a fluid flow in a first axial direction, a circumferential fluid channel configured to direct the fluid flow in a circumferential direction, and a second fluid channel configured to direct the fluid flow in a second axial direction substantially opposite the first axial direction. The first fluid channel includes a first outlet passing through the inner surface, and the second fluid channel including a second outlet passing through the inner surface.