XFOIL Version 6.94 Calculated polar for: coude3 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.022 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.3988 0.08508 0.07808 0.0351 0.9999 0.5169 -2.750 -0.3402 0.07922 0.07192 0.0152 0.9999 0.4428 -2.500 -0.2847 0.07489 0.06730 0.0018 0.9999 0.3975 -2.250 -0.2381 0.07150 0.06375 -0.0060 0.9999 0.3690 -2.000 -0.1905 0.06869 0.06077 -0.0131 0.9999 0.3449 -1.750 -0.1438 0.06643 0.05840 -0.0187 0.9999 0.3266 -1.500 -0.0920 0.06467 0.05660 -0.0246 0.9999 0.3109 -1.250 0.1448 0.05318 0.04280 -0.0631 0.5737 0.3049 -1.000 0.2177 0.05277 0.04107 -0.0710 0.5189 0.3060 -0.750 0.2953 0.05264 0.03995 -0.0799 0.4860 0.3097 -0.500 0.3757 0.05257 0.03950 -0.0896 0.4663 0.3254 -0.250 0.4449 0.05267 0.03972 -0.0974 0.4535 0.3532 0.000 0.5151 0.05284 0.04039 -0.1055 0.4412 0.4083 0.250 0.5719 0.05164 0.04083 -0.1093 0.4299 1.0001 0.500 0.6489 0.05429 0.04212 -0.1177 0.4187 1.0001 0.750 0.7056 0.05649 0.04388 -0.1234 0.4156 1.0001 1.000 0.7592 0.05896 0.04613 -0.1291 0.4147 1.0001 1.250 0.8046 0.06154 0.04864 -0.1334 0.4155 1.0001 1.500 0.8433 0.06428 0.05136 -0.1365 0.4176 1.0001 1.750 0.8761 0.06724 0.05430 -0.1384 0.4203 1.0001 2.000 0.8841 0.06953 0.05672 -0.1352 0.4232 1.0001 2.250 0.8491 0.07099 0.05856 -0.1238 0.4288 1.0001 2.500 0.8295 0.07353 0.06130 -0.1162 0.4340 1.0001 2.750 0.8151 0.07653 0.06441 -0.1099 0.4388 1.0001 3.000 0.8096 0.07995 0.06785 -0.1056 0.4426 1.0001 3.500 0.6632 0.08916 0.07790 -0.0822 0.4726 1.0001 3.750 0.6218 0.09536 0.08428 -0.0789 0.4920 1.0001 4.000 0.5128 0.10579 0.09538 -0.0775 0.5491 1.0001 4.750 0.2089 0.12383 0.11608 -0.0618 0.9691 1.0001 5.000 0.2181 0.12624 0.11804 -0.0613 0.9722 1.0001