A turbine blade in a gas turbine engine includes an airfoil extending in a radial direction. The airfoil has an outer wall delimiting an airfoil interior. The outer wall includes a pressure sidewall and a suction sidewall joined at a leading edge and a trailing edge in a longitudinal direction. A turbulator is disposed in the airfoil interior. The turbulator includes a first row having at least two turbulator ribs spaced apart in the longitudinal direction. The turbulator includes a second row extending in the radial direction from the first row and having at least two turbulator ribs spaced apart in the longitudinal direction.

A turbine blade for a gas turbine engine has: an airfoil extending along a span from a base to a tip and along a chord from a leading edge to a trailing edge, the airfoil having a pressure side and a suction side, a tip pocket at the tip of the airfoil, the tip pocket at least partially surrounded by a peripheral tip wall defining a portion of the pressure and suction sides; at least one internal cooling passage in the airfoil and having at least one outlet communicating with the tip pocket; and a reinforcing bump located on the pressure side of the airfoil and protruding from a baseline surface of the peripheral tip wall to a bump end located into the tip pocket, the reinforcing bump overlapping a location where a curvature of a concave portion of the pressure side of the airfoil is maximal.

The air moving device includes a rotor and a stator. The quantity of the rotor blades is not less than 5 and not greater than 12. The average blade angle of rotor blades is not less than 45 degrees and is not greater than 64 degrees. The ratio of the hub diameter to the rotor diameter is not less than 0.4 and not greater than 0.79. The quantity of the stator blades is not less than 6 and not greater than 23. The average blade angle of stator blades is not less than 45 degrees and not greater than 70 degrees. The ratio of the total thickness of the air moving device to the rotor diameter is not less than 0.76 and not greater than 1.7. The ratio of the stator axial thickness to the rotor axial thickness is not less than 0.28 and not greater than 0.65.

The impeller of a rotating machine according to at least one embodiment of the present discloser is provided with: a disc; a cover disposed on an opposite side of a radial passage from the disc in an axial direction; and a blade disposed between the disc and the cover. In a dimensionless position along a camber line of the blade when the position of a leading edge of the blade is defined as 0 and the position of a trailing edge of the blade is defined as 1, a position where an angle difference between a first blade angle at a disc-side end portion of the blade and a second blade angle at a cover-side end portion of the blade is maximum is in a range of 0.5 or more and 1 or less. The first blade angle is −10 degrees or more and 0 degrees or less at the position where the angle difference is maximum.

Core assemblies for manufacturing airfoils and airfoils made therefrom are described. The core assemblies having a leading edge hybrid skin core positioned relative to a plurality of core bodies and configured to define a leading edge cavity at a leading edge of the manufactured airfoil. The leading edge hybrid skin core extends from a root region toward a tip region in a radial direction, the leading edge hybrid skin core extends above at least one of the plurality of core bodies to define an exit in a tip region of the manufactured airfoil, and the leading edge hybrid skin core has a height-to-width ratio of about 0.8 or less.

A rotor blade includes an airfoil having an airfoil shape. The airfoil shape has a nominal profile substantially in accordance with Cartesian coordinate values of X, Y and Z set forth in one of Table I, Table II, Table III, Table IV, Table V, Table VI, Table VII, Table VIII, or Table IX. The Cartesian coordinate values of X, Y and Z are non-dimensional values from 0% to 100% convertible to dimensional distances expressed in a unit of distance by multiplying the Cartesian coordinate values of X, Y and Z by a scaling factor of the airfoil in the unit of distance. The X and Y values, when connected by smooth continuing arcs, define airfoil profile sections at each Z value. The airfoil profile sections at Z values are joined smoothly with one another to form a complete airfoil shape.

An airfoil for a gas turbine engine includes a first portion joined to a second portion along an interface such that at least the first portion establishes an airfoil section and the second portion establishes a root section. The airfoil section includes an airfoil body that extends between leading and trailing edges in a chordwise direction, between pressure and suction sides separated in a thickness direction, and from the root section to a tip portion in a spanwise direction. A recessed region extends inwardly from at least one of the pressure and suction sides. The airfoil body includes at least one rib that bounds a respective pocket within a perimeter of the recessed region. The recessed region and the at least one rib are dimensioned to extend across the interface. A cover skin is coupled to the airfoil body along the at least one rib to enclose the recessed region.

A gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, a propulsor section, a geared architecture, a high spool and a low spool. The high spool includes a high pressure compressor and a high pressure turbine. The low spool includes a low pressure compressor and a low pressure turbine. At least one stage of the turbine section includes an array of rotatable blades and an array of vanes. A ratio of the number of vanes to the number blades is greater than or equal to 1.55. A mechanical tip rotational Mach number of the blades is greater than or equal to 0.5 at an approach speed.

An airfoil assembly for a gas turbine engine includes a vane, a support strut, and an inner carrier. The vane is adapted to interact with hot gases flowing around the airfoil assembly during use of the airfoil assembly. The support strut is located in an interior cavity formed in the vane and configured to receive force loads applied to the ceramic matrix composite vane by the hot gases. The inner carrier is coupled with the support strut and adapted to block the hot gases from flowing radially inward toward an axis of the gas turbine engine.

A ceiling fan assembly including a non-rotating motor shaft with an upper and lower bearing stop, a stator mounted to the non-rotating motor shaft, a rotor surrounding the stator, a motor housing having an upper bearing seat spaced above the upper bearing stop and a lower bearing seat spaced below the lower bearing stop, an upper bearing seated within the upper bearing seat, a lower bearing seated within the lower bearing seat, and a downrod coupling provided on the non-rotating shaft. When the ceiling fan assembly is suspended from a structure with the downrod coupling, the weight of the rotor presses the upper bearing against the upper bearing stop such that the weight of the rotor is transferred through the upper bearing to the non-rotating shaft.