An Improved Blade Root Design for Axial Flow Compressors (and Turbines)

Abstract

To save weight, designers of aircraft gas turbines aim to produce axial compressors with fewer rows of blades and greater mass flows per unit of frontal area. These improvements often consist of increases in blade tip speeds and reductions in the ratio of the hub to tip diameters. As the efficiency of transonic blading is partly dependent upon a low value of blade pitch/chord it is now frequently necessary to design a blade with a greater actual chord at the tip than at the hub.