A turbine blade is provided. The turbine blade may include a blade extending from a platform to a free end and having an airfoil-shaped cross section, the blade including a leading edge, a trailing edge, a pressure side extending from the leading edge to the trailing edge, and a suction side extending from the leading edge to the trailing edge, one or more internal cooling passages through which cooling air flows, a trailing edge slot formed along the trailing edge and connected to the internal cooling passage, and a pin-fin array including a plurality of pin-fins positioned in the internal cooling passage connected to the trailing edge slot, each pin-fin including a main body and chamfered or filleted portions respectively connected to the pressure side and the suction side at respective ends of the main body, wherein among the pin-fins of the pin-fin array, a portion of the pin-fins have relatively large chamfered or filleted portions as compared with remaining pin-fins.

Turbomachinery hardware, used in a rotor assembly and a stator assembly, including an airfoil portion including a leading edge, a trailing edge, a pressure side, and a suction side, and a platform on which the airfoil portion is disposed. The platform including a platform axis, a pressure side mateface located adjacent to the pressure side of the airfoil portion and a suction side mateface located adjacent to the suction side airfoil portion, wherein a portion of a pressure side mateface includes a first geometry, and a portion of a suction side mateface includes a second geometry. The first geometry is selected from a group consisting of: oblique to a platform axis, and a first curved portion. The second geometry is selected from a group consisting of: oblique to the platform axis and a second curved portion.

A turbine vane includes an airfoil having a leading edge and a trailing edge, an inner shroud disposed at one end of the airfoil to support the airfoil, and an outer shroud disposed opposite to the inner shroud at the other end of the airfoil to support the airfoil, wherein a corner part is formed at a portion where the airfoil and the inner shroud or the outer shroud meet, the corner part including a first round portion connected in an arc shape to the inner shroud or the outer shroud, a first inclined portion connected to the first round portion and outwardly extending in an inclined shape, and a second round portion connected to the first inclined portion and outwardly extending in an arc shape.

An internal cooling system (14) including an impingement jet strike channel system (16) for increasing the effectiveness of impingement jets (18) is disclosed. The impingement jet strike channel system (16) may include an impingement jet strike cavity (20) offset from one or more impingement orifices (22). A plurality of impingement jet strike channels (24) may extend radially outward from the impingement jet strike cavity (20) forming a starburst pattern of impingement jet strike channels (24) and may be formed by a plurality of ribs (26) that each separate adjacent impingement jet strike channels (24). The ribs (26) forming the impingement jet strike channels (24) may be split one or more times into multiple channels to increase the number of stagnation points (28, 38, 52) to increase the cooling capacity. The impingement jet strike channel system (16) may be used within components, such as, but not limited to, gas turbine engines (12), including vane inserts, airfoil leading edge cooling systems, platforms, advanced transitions, acoustic resonators, ring segments and the like.

A gas turbine engine 10 is provided in which a fan having fan blades 139 in which the camber distribution relative to covered passage of the fan 13 allows the gas turbine engine to operate with improved efficiency when compared with conventional engines, whilst retaining an acceptable flutter margin.

A gas turbine engine 10 is provided in which a fan having fan blades 139 in which the camber distribution relative to covered passage of the fan 13 allows the gas turbine engine to operate with improved efficiency when compared with conventional engines, whilst retaining an acceptable flutter margin.

Turbine nozzles are provided for gas turbine engines. The turbine nozzle includes an arcuate inner band; an arcuate outer band; and a nozzle vane disposed between the arcuate inner band and the arcuate outer band. The radially inner end of the nozzle vane is attached to the arcuate inner band through an interlocking transition zone including a plurality of projections alternately extending from the radially inner end of the nozzle vane and the arcuate inner band, respectively, to undetachably couple the nozzle vane and the arcuate inner band. Optionally, the radially outer end of each nozzle vane is also attached to the arcuate outer band through an interlocking transition zone.

A component for a gas turbine engine according to an exemplary aspect of this disclosure includes a surface having a plurality of cooling holes. The surface includes a first region and a second region each having a plurality of cooling holes. The cooling holes within the first region are arranged differently than the cooling holes in the second region.

An internal cooling system (14) including an impingement jet strike channel system (16) for increasing the effectiveness of impingement jets (18) is disclosed. The impingement jet strike channel system (16) may include an impingement jet strike cavity (20) offset from one or more impingement orifices (22). A plurality of impingement jet strike channels (24) may extend radially outward from the impingement jet strike cavity (20) forming a starburst pattern of impingement jet strike channels (24) and may be formed by a plurality of ribs (26) that each separate adjacent impingement jet strike channels (24). The ribs (26) forming the impingement jet strike channels (24) may be split one or more times into multiple channels to increase the number of stagnation points (28, 38, 52) to increase the cooling capacity. The impingement jet strike channel system (16) may be used within components, such as, but not limited to, gas turbine engines (12), including vane inserts, airfoil leading edge cooling systems, platforms, advanced transitions, acoustic resonators, ring segments and the like.

An axial fan includes an impeller that is provided with a blade that rotates around a central axis extending in a vertical direction, a motor unit that rotates the impeller, and a housing that is provided closer to a radially outer side than the impeller and surrounds the impeller. The impeller is further provided with an assist blade that protrudes at least in an axial direction from a radially outer end portion of the blade and extends in a circumferential direction. The assist blade is provided with a variable width portion at which an assist blade width between the blade and a tip end portion of the assist blade in a direction in which the assist blade protrudes decreases toward a rear side in a rotation direction from a front side in the rotation direction.