Performance Study of Novel Compressor Blades in a Two-Dimensional Cascade—Transonic Regime

Passengers would always like to reach their destinations with minimum commute time. Generating a higher thrust is a necessity. This implies that the turbomachinery associated with the power plant has to rotate faster and with higher efficiencies. However, high rotational speeds, mainly in the transonic regime, often lead to boundary layer separation, shocks, compressor stall, and surge. The current investigation is an attempt to reduce the abovementioned phenomena. It involves the performance study of a smoothened controlled diffusion airfoil (CDA) blade that has been optimized by “Multi-Objective Genetic Algorithm” (MOGA) by altering maximum camber location and stagger angle. Inlet pressure is varied from 15 kPa to 30 kPa and the angle of attack ranging from 40.4° to 56.4°. C48-S16-BS1 is validated and considered as the baseline profile, and all other blades are collated to this. It is observed that shifting the location of the maximum camber close to the leading edge and increasing stagger angle result in improvement of blade performance in terms of lower pressure losses for high angles of attack. Shifting the camber location slightly lesser than the mid-chord and increasing the stagger angle showed the best performance throughout. However, moving the camber location close to the trailing edge always resulted in the highest amount of losses due to its poor performance. Furthermore, a higher stagger angle is preferred.