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 = 0.035 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.1805 0.02792 0.01556 -0.0320 1.0000 0.1069 -2.750 -0.1507 0.02615 0.01343 -0.0309 1.0000 0.1066 -2.500 -0.1227 0.02460 0.01163 -0.0295 1.0000 0.1086 -2.250 -0.0961 0.02330 0.01015 -0.0281 1.0000 0.1138 -2.000 -0.0699 0.02198 0.00893 -0.0270 1.0000 0.1271 -1.750 -0.0425 0.02018 0.00779 -0.0266 1.0000 0.1977 -1.000 0.0285 0.01728 0.00655 -0.0209 1.0000 1.0000 -0.750 0.0536 0.01736 0.00648 -0.0198 1.0000 1.0000 -0.500 0.0788 0.01744 0.00651 -0.0188 1.0000 1.0000 -0.250 0.1045 0.01753 0.00666 -0.0176 1.0000 1.0000 0.000 0.1324 0.01758 0.00688 -0.0163 1.0000 1.0000 0.250 0.2087 0.02118 0.00749 -0.0223 0.2866 1.0000 0.500 0.2331 0.02233 0.00821 -0.0213 0.2617 1.0000 0.750 0.2589 0.02346 0.00900 -0.0205 0.2408 1.0000 1.000 0.2867 0.02457 0.00992 -0.0199 0.2279 1.0000 1.250 0.3172 0.02567 0.01104 -0.0196 0.2230 1.0000 1.500 0.3491 0.02684 0.01233 -0.0195 0.2199 1.0000 1.750 0.3819 0.02813 0.01383 -0.0196 0.2188 1.0000 2.000 0.4136 0.02956 0.01552 -0.0195 0.2195 1.0000 2.250 0.4449 0.03121 0.01750 -0.0196 0.2216 1.0000 2.500 0.4750 0.03313 0.01965 -0.0197 0.2248 1.0000 2.750 0.5061 0.03498 0.02198 -0.0200 0.2308 1.0000 3.000 0.5370 0.03735 0.02479 -0.0206 0.2395 1.0000 3.250 0.5690 0.03985 0.02787 -0.0217 0.2526 1.0000 3.500 0.5965 0.04295 0.03113 -0.0221 0.2593 1.0000 3.750 0.6276 0.04575 0.03466 -0.0243 0.2713 1.0000 4.000 0.6577 0.04941 0.03877 -0.0266 0.2906 1.0000 4.250 0.7084 0.05511 0.04574 -0.0409 0.3890 1.0000 4.750 0.6317 0.07036 0.06144 -0.0885 0.7374 1.0000 5.000 0.5106 0.06882 0.05956 -0.0778 0.9336 1.0000