A compressor airfoil (12) includes pressure and suction sides (18,20) extending from root (22) to tip (24) and between leading and trailing edges (26,28). Transverse sections have respective chords and camber lines. Centers of gravity (34) of the sections are aligned along a bowed stacking axis either tangentially, axially, or both, for improving performance.

Bowed compressor airfoil

A turbine nozzle includes outer and inner bands between which extend a plurality of vanes for channeling combustion gases. Each of the vanes includes leading and trailing edges, and pressure and suction sides extending therebetween, and also a bow along the trailing edge to increase pressure in the gases adjacent the inner band. The vanes also include a thermal barrier coating (TBC) selectively disposed solely along the suction side between the leading and trailing edges.

Bowed nozzle vane with selective TBC

A turbine stator vane (41) for use in an axial flow gas turbine. The vane has an aerofoil the pressure face of which is convex between platform (45) and tip (46) regions in a plane (48) which extends both radially of the turbine and transversely of the general working fluid flow direction between the vanes. The trailing edge (43) of the aerofoil is straight from platform to tip, and the spanwise convex and concave curvatures of the aerofoil pressure and suction surfaces respectively are achieved by rotational displacement of the aerofoil sections about the straight trailing edge. However, the axial width (W) of the aerofoil is substantially constant over substantially all of the aerofoil radial height and the chord line at mid-height aerofoil sections (44) is shorter than the chord lines in aerofoil sections at platform or tip regions. Reducing chord length at the mid-height region in this way lowers aerodynamic profile losses without unduly affecting vane performance. Also disclosed is a turbine rotor blade designed to form a stage pair with the stator vane.

Aerofoil for an axial flow turbomachine

An airfoil for a turbine engine includes an airfoil having 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 a leading edge dihedral and a span position. The leading edge dihedral is negative from the 0% span position to the 100% span position. A positive dihedral corresponds to suction side-leaning, and a negative dihedral corresponds to pressure side-leaning.

Gas turbine engine airfoil

The invention relates to a turbine engine blade, the airfoil of which extends radially between a blade root and an airfoil tip, axially between a leading edge (2) and a trailing edge (3), and tangentially between a pressure side and a suction side (5), the profile of said blade consisting of a series of basic profiles, in the form of a vane section, stacked on one another along a so-called stacking line connecting the centre of gravity of all the vane sections, characterised in that the projection of said stacking line of the airfoil on at least one plane extending radially from the blade root comprises a double tangential inversion of the direction of the curvature thereof, located in the last thirty percent of the height of the airfoil, the projection plane being positioned substantially perpendicular to the chord of the blade.

Turbine engine blade having improved stacking law

A rotor blade (24) for a gas turbine engine includes a bowed surface (43) on a tip region (40) of the suction side (32) thereof. The curvature of the bowed surface (43) progressively increases toward the tip (36) of the blade (24). The bowed surface (43) results in a reduction of tip leakage through a tip clearance (50) from the pressure side (30) to the suction side (32) of the blade (24) and reduces mixing loss due to tip leakage.

Joseph Staubach

Papers

Rotor airfoils to control tip leakage flows

Rotorschaufeln zur verminderung von spaltverlusten