A rotor blade for a gas turbine engine includes a structural member formed from a metallic material. The structural member, in turn, includes a base portion, a spar, and a tip cap, with the base portion at least partially forming a root of the rotor blade and a shank of the rotor blade. Furthermore, the structural member includes a fairing formed from a composite material. The fairing is, in turn, coupled to the structural member such that the fairing forms at least a portion of an airfoil of the rotor blade and a platform of the rotor blade. Additionally, the fairing including a first fairing panel and a second fairing panel in contact with the first fairing panel at a first split line and a second split line.

An impeller includes a hub and blades. In radial direction, an upper surface of the blade near the outer edge includes a groove structure, a lower surface of the blade relating to the groove structure includes a peak structure, the lower surface of the blade includes a recess structure aside the trailing edge or the leading edge. The recess structure is connected to the peak structure. The blade has at least five airfoils from the inner edge to the outer edge. The blade is defined by continuously connecting the at least five airfoils at different sections in sequence so as to form the groove structure and the peak structure respectively on the upper surface and the lower surface. The groove structure and the peak structure are aside the outer edge of the blade.

An impeller associated with a compressor of a gas turbine engine includes a plurality of impeller blades. Each impeller blade of the plurality of impeller blades has a leading edge and a trailing edge opposite the leading edge in a streamwise direction. Each impeller blade of the plurality of impeller blades extends in a spanwise direction from a hub at 0% span to a tip at 100% span, and each impeller blade of the plurality of impeller blades has a plurality of mean camber lines that each extend from the leading edge to the trailing edge at a respective spanwise location. The leading edge includes a partially swept leading edge surface defined between 70% span to 100% span that extends in the streamwise direction between 3% to 15% of a mean camber line at 100% span.

Various embodiments of the disclosure include turbine blades and systems employing such blades. Various embodiments include a turbine blade having: an airfoil having an airfoil shape having a nominal profile substantially in accordance with at least a portion of Cartesian coordinate values of X, Y and Z set forth in TABLE I. The Cartesian coordinate values are non-dimensional values of from 0% to 100% convertible to distances by multiplying the values by a height of the airfoil expressed in units of distance. The X and Y values are connected by smooth continuing arcs to define airfoil profile sections at each distance Z along at least a portion of the airfoil, the profile sections at the Z distances being joined smoothly with one another to form the nominal profile.

A blade of a steam turbine includes a plurality of turbine blade rows which are fixed to a radially outer side of a rotor shaft rotating about an axis, and are arranged in an axial direction in which the axis extends, and a turbine vane row which is disposed to be adjacent to an upstream side of the turbine blade row in the axial direction for each of the plurality of turbine blade rows, the blade of a steam turbine including a blade body which is disposed in a steam main flow path which is formed around a rotary shaft such that main steam flows through the steam main flow path, the blade body having an airfoil cross section in which a concave positive-pressure surface and a convex negative-pressure surface are continuous to each other via a leading edge and a trailing edge.

A blade of a fan or compressor that reduces loss by enlarging a laminar flow region over a blade surface is provided. The blade is divided into a subsonic region where the relative Mach number of the inlet air flow during rated operation of a turbofan engine is lower than 0.8 and a transonic region where the relative Mach number is equal to or higher than 0.8. A blade surface angle change rate is based on an angle formed by a tangent to the blade surface and the axis of the engine, the leading edge blade surface angle, and the trailing edge blade surface angle at. In each of the subsonic region and the transonic region, values of the blade surface angle change rate on the pressure and suction surfaces are defined at predetermined axial locations along the chord on the pressure and suction surfaces.

A blade, an impeller and a fan are provided. A trailing edge (12) of the blade (1) is provided with at least one concave arc segment (15). At least one end point of the at least one concave arc segment (15) is located between a radial outer edge (13) of the blade (1) and a radial inner edge (14) of the blade (1). The blade (1) is provided with at least one ridge structure protruding from a pressure surface (18) of the blade (1) toward a suction surface (17) of the blade (1). The blade is provided with at least one concave arc segment at a trailing edge thereof on the basis of a bionics principle, and is further provided with a ridge structure, thus improving an airflow pattern at the trailing edge of the blade by means of changing a shape of the blade, and reducing noise accordingly.

An apparatus and method regarding an airfoil for a turbine engine, the airfoil comprising an outer wall defining an interior bound by a pressure side and a suction side extending axially between a leading edge and a trailing edge defining a chord-wise direction and extending radially between a root and a tip defining a span-wise direction. The airfoil further comprising at least one cooling passage extending radially within the interior and defining a primary cooling airflow; and at least one cooling hole having an inlet defining a first cross-sectional area, the inlet in communication with the cooling passage, an outlet defining a second cross-sectional area greater than the first cross-sectional area; wherein the primary cooling airflow enters the inlet in a first direction and exits the outlet in a second direction different than the first direction.

A centrifugal blower includes a centrifugal fan and a case that houses the centrifugal fan. A blade of the centrifugal fan has an intake side edge portion extending from a shroud inward in a radial direction. The intake side edge portion has a negative pressure side inclined portion inclined to a negative pressure surface side of the blade with respect to an axial direction. The negative pressure side inclined portion has an inclination length in the axial direction, and the inclination length is larger in a reverse flow portion of the intake side edge portion close to the shroud than in a radially innermost portion of the intake side edge portion located at an innermost side in the radial direction.

Airfoil assemblies for gas turbine engines are provided. For example, an airfoil assembly comprises an airfoil, an inner band defining an inner opening shaped complementary to an inner end of the airfoil, and an outer band defining an outer opening shaped complementary to an outer end of the airfoil. The airfoil inner end is received with the inner opening, and the airfoil outer end is received within the outer opening. A strut extends radially through an airfoil cavity. A first pad is defined at a first radial location within the cavity. A second pad is defined within the cavity at a second, different radial location. In some embodiments, the airfoil assembly inner band includes a first inner flange, through which the inner band is secured to a support structure, and the outer band includes a first outer flange, through which the outer band is secured to a support structure.