In many industrial gas turbine compressor designs, the compressors later stage blade angles are reduced in a constant flow area section as a means to even out the per stage workload. Most compressors use NACA 65 series type airfoils, which are good for high subsonic and supersonic flow, but are poor for middle or low subsonic flows. The temperature increases as the compression ratio increases; which cause the Mach number to drop. With reduced blade cascade overlaps, a reduction in axial blade solidity results. The compounding effect of low solidity and a low Mach number can cut the stalling angle by several degrees. This recent study found that compressor stall more or less is linked to the change of moment coefficient Cm, rather than lift coefficient Cl. Designing the airfoil, by extending the constant moment coefficient to a higher angle of attack region can delay the trailing edge upper surface separation to a higher angle of attack, the main source of rotating stall. This separated flow exhibits itself more clearly on the moment coefficient, but is obscured by an increase in lift coefficient before “aerodynamic” stall. This new design is based on the second order derivative of the camber line, with a low drag symmetrical airfoil thickness. Numerical simulation of a single airfoil and cascade of the new airfoil is compared with other shapes. The results show that the trailing edge flow separation begins at a 9.5-degree angle of attack for the NACA 65 series airfoils. The NACA 0012 separation (i.e. change in Cm) starts at 5 degrees (total stall occurs at 11 degrees). The new airfoil CFS18-0010 exhibits no separation for a single airfoil of up to 12 degrees. The cascade results showed no flow separation up to an angle of 15 degrees, which is enough to eliminate most of the rotating stall.

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