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.

A vane includes an airfoil that defines a leading edge, a trailing edge, a pressure side, and a suction side. The airfoil includes a truss structure that has ribs that define there between a plurality of through-cavities from the pressure side to the suction side. Honeycomb cells are disposed in the cavities. A face sheet defines at least one of the pressure side or the suction side. The face sheet has perforations that correspond in location to the honeycomb cells.

A fan includes a first wind wheel having a first rotation axis and a second wind wheel arranged opposite to the first wind wheel and having a second rotation axis. The first wind wheel includes a first hub and a plurality of first blades arranged around the first hub and spaced apart from one another. The second wind wheel includes a second hub not directly coupled to the first hub and a plurality of second blades arranged around the second hub and spaced apart from one another. The first blades and the second blades are tilted towards opposite directions along a circumferential direction.

Disclosed is a method of detecting mixed mode corrosion of a turbine airfoil. The method includes heat treating the turbine airfoil at a temperature of 1600 to 2100° F. in a Ni/Co oxide reducing atmosphere; and scanning the heat treated turbine airfoil with a magnetometer to determine the presence of mixed mode corrosion, wherein the turbine airfoil comprises a nickel superalloy and has internal passages.

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 gas turbine engine includes a compressor and a turbine; a plurality of rotor blades positioned in the compressor or the turbine and arranged in a plurality of stages along an axial direction; an outer drum positioned at least partially radially outward of the plurality of rotor blades and rotatable with the plurality of rotor blades; and a casing extending over the outer drum and defining a plurality of air cavities with the outer drum, the plurality of air cavities arranged in series airflow communication along the axial direction.

An airfoil includes an airfoil wall that defines a leading end, a trailing end, a first side wall, and a second side wall. A rib connects the first and second side walls of the airfoil wall. The rib includes first and second unbranched arms that initiate at the second side wall and extend there from to meet at a node. A third unbranched arm initiates at the first side wall. The third unbranched arm extends perpendicularly there from and connects at either the node, the first unbranched arm, or the unbranched second arm.

A gas turbine engine 10 is provided in a fan root to tip pressure ratio, defined as the ratio of the mean total pressure of the flow at the fan exit that subsequently flows through the engine core (P.sub.102) to the mean total pressure of the flow at the fan exit that subsequently flows through the bypass duct (P.sub.104), is no greater than a certain value. The gas turbine engine 10 may provide improved efficiency when compared with conventional engines, whilst retaining an acceptable flutter margin.

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 gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, a turbine section including a fan drive turbine, a geared architecture driven by the fan drive turbine, and a fan driven by the fan drive turbine via the geared architecture. 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 about 1.55. A mechanical tip rotational Mach number of the blades is configured to be greater than or equal to about 0.5 at an approach speed.