XFOIL Version 6.94 Calculated polar for: GOA 1 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.008 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.1906 0.03617 0.01821 -0.0236 1.0000 1.0000 -2.750 -0.1577 0.03601 0.01588 -0.0236 1.0000 1.0000 -2.500 -0.1310 0.03586 0.01445 -0.0227 1.0000 1.0000 -2.250 -0.1055 0.03574 0.01326 -0.0218 1.0000 1.0000 -2.000 -0.0806 0.03565 0.01237 -0.0210 1.0000 1.0000 -1.750 -0.0560 0.03559 0.01163 -0.0202 1.0000 1.0000 -1.500 -0.0317 0.03555 0.01103 -0.0194 1.0000 1.0000 -1.250 -0.0074 0.03554 0.01056 -0.0186 1.0000 1.0000 -1.000 0.0166 0.03555 0.01021 -0.0178 1.0000 1.0000 -0.750 0.0406 0.03558 0.00998 -0.0170 1.0000 1.0000 -0.500 0.0645 0.03564 0.00979 -0.0162 1.0000 1.0000 -0.250 0.0884 0.03572 0.00971 -0.0154 1.0000 1.0000 0.000 0.1126 0.03581 0.00952 -0.0146 1.0000 1.0000 0.250 0.1364 0.03594 0.00969 -0.0137 1.0000 1.0000 0.500 0.1603 0.03609 0.00999 -0.0129 1.0000 1.0000 0.750 0.1843 0.03626 0.01042 -0.0121 1.0000 1.0000 1.000 0.2087 0.03645 0.01102 -0.0112 1.0000 1.0000 1.250 0.2338 0.03665 0.01181 -0.0105 1.0000 1.0000 1.500 0.2605 0.03684 0.01286 -0.0098 1.0000 1.0000 1.750 0.2908 0.03703 0.01429 -0.0098 1.0000 1.0000 2.000 0.3254 0.03739 0.01637 -0.0119 1.0000 1.0000 2.250 0.3451 0.03893 0.01921 -0.0162 1.0000 1.0000 2.500 0.3428 0.04196 0.02222 -0.0197 1.0000 1.0000 2.750 0.3424 0.04509 0.02508 -0.0230 1.0000 1.0000 3.000 0.5241 0.05385 0.03462 -0.0587 0.7759 1.0000 3.250 0.5684 0.05719 0.03801 -0.0629 0.7312 1.0000 3.500 0.5912 0.06061 0.04140 -0.0654 0.7145 1.0000 3.750 0.6008 0.06404 0.04477 -0.0671 0.7081 1.0000 4.000 0.6124 0.06761 0.04828 -0.0689 0.7039 1.0000 4.250 0.6201 0.07117 0.05177 -0.0705 0.7032 1.0000 4.500 0.6223 0.07459 0.05508 -0.0714 0.7053 1.0000 4.750 0.6254 0.07809 0.05848 -0.0725 0.7091 1.0000 5.000 0.6326 0.08182 0.06214 -0.0742 0.7139 1.0000