CFD Analysis of Axial Compressor Rotor with Leading-Edge Tubercles

Abstract

An axial compressor is a fundamental part of a jet engine. The efficiency of a jet engine heavily depends upon the efficiency of its compressor. In this present study, CFD analysis has been undertaken to determine the influence of sinusoidal leading-edge tubercles on transonic axial compressor efficiency. Although several studies on the effect of tubercles on various isolated blades are available in literature, a detailed study of their effect on axial compressor rotor blade was lacking. Presently analysis has been performed on NASA Rotor 37. Many cases were considered by changing different tubercles parameters: amplitude, wavelength, span wise location of tubercles and height of tubercles (amplitude in y-direction) on rotor blade. The results predict an increase of 0.57% in the efficiency of the modified blade in comparison with the corresponding baseline blade which is maintained for entire mass flow rate operating range. The increase in efficiency is due to higher pressure at outlet in modified cases. Higher value of pressure at outlet is attributed to change in shockwave pattern for all the modified cases. The shockwave has been modified from normal biased shockwave to oblique-biased shockwave. Oblique shockwave being less detrimental is well documented due to reduction in aerodynamic loss. Finally, alteration in separation is observed near the suction side of tubercle model near mid of chord. This change in separation pattern carries high energy flow near the blade surface, reenergizing the boundary layer.