A radial compressor impeller having a wheel disc, a cover disc arranged at a distance from the wheel disc, and blades arranged therebetween and connecting the wheel disc and the cover disc. The blades have an end portion, which is arranged at a distance from the radial outer end of the wheel disc and the cover disc towards the inside and has a reduced thickness with respect to an adjoining section.

A fan assembly for a gas turbine engine. The assembly includes a hub and a plurality of blades projecting from the hub. The hub includes an inlet aperture located adjacent a blade root. The inlet aperture is in fluid communication with a passage extending along at least part of a span of the blade, and communicates with an outlet provided on an exterior of the blade, radially outwardly of the inlet.

A centrifugal fan impeller structure includes a hub and a blade body set. The hub has an extension section. The blade body set has multiple blade bodies. The blade bodies outward extend from the extension section of the hub. Each two adjacent blade bodies define therebetween a flow way, an air outlet and an air inlet. The air outlet and the air inlet are respectively positioned at two ends of the flow way in communication with the flow way. The air outlets are arranged at unequal intervals so as to greatly reduce noise in operation.

A ceiling mounted fan or ceiling fan includes a body defining an interior passage having an inlet and an outlet provided on the body. The inlet, outlet, and interior passage can be annular. An impeller mounted is mounted within the interior passage and driven by a motor mounted within the body to draw a volume of air through the interior passage from the inlet to the outlet. The impeller blades include a set of extensions extending from the tip.

A turbine blade includes cooling passages formed inside the blade and extending in a blade height direction, blade surfaces on a suction side and a pressure side being covered with thermal barrier coating, a design point on a suction side being set on the blade surface on the suction side of each blade section perpendicular to the blade height direction within a range from a position on a back side of and including a throat position, and to a position in front of and not including a tailing end of a final cooling passage. Thickness distribution of the thermal barrier coating on the suction side of each blade section is configured such that a thickness of the thermal barrier coating is uniform from a blade leading edge to the design point and gradually reduces from the design point toward the back side up to the blade trailing edge.

An airfoil of a gas turbine engine includes an airfoil body having a leading edge and a trailing edge extending in a radial direction, a trailing edge cavity formed within the airfoil and proximate to the trailing edge of the airfoil, the trailing edge cavity extending from the trailing edge in a forward direction toward the leading edge, at least one set of blocking pedestals located within the trailing edge cavity, a set of circular pedestals located aftward from the at least one blocking set of pedestals, and a set of spear pedestals located aftward from the set of circular pedestals and closest to the trailing edge of the airfoil body.

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.

An airfoil for a turbine engine includes pressure and suction sides extending in a radial direction from a 0% span position at an inner flow path location to a 100% span position at an airfoil tip. The airfoil has a relationship between an axial leading edge location and a span position that is at least a third order polynomial with a generally U-shaped curve having an initial negative slope followed by a positive slope. The positive slope leans aftward and the negative slope leans forward. The curve has a critical point in the range of 30-50% span position at which the curve changes from the negative slope to the positive slope. The curve is generally linear from 55% span to 75% span and has a positive slope that increases at a rate of about 0.0875 inch (2.22 mm) per 1% span, +/−0.04 inch (1.01 mm) per 1% span.

Profiled blade for a fan wheel, with a profile body produced from at least one curved metal strip, whereby adjacent end zones of the at least one metal strip are connected to each other in an adhesively bonded and/or form-fit manner wherein a second end zone at a distance from a second end zone edge is equipped with at least one embossed area for forming a contact surface enabling a first end zone to, at least in sections, be brought into contact with it.

A turbine airfoil for a gas turbine engine includes a pressure sidewall extending along a spanwise direction, and from a leading edge of the airfoil towards the trailing edge of the airfoil. The turbine airfoil additionally includes a suction sidewall also extending along the spanwise direction, and from the leading edge towards the trailing edge. The pressure sidewall and suction sidewall define a cooling air cavity therebetween, and one or both of the pressure sidewall and suction sidewall define a trailing edge cooling channel extending from the cooling air cavity substantially to the trailing edge. Additionally, one or both of the pressure sidewall and suction sidewall include a plurality of pressure drop members extending partially into the trailing edge cooling channel for reducing an amount of cooling air flowing therethrough from the cooling air cavity.