XFOIL Version 6.94 Calculated polar for: manu4/33 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.018 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.4564 0.07358 0.06509 0.0577 0.9999 0.7005 -2.750 -0.4439 0.07188 0.06331 0.0514 0.9999 0.6548 -2.500 -0.4151 0.07010 0.06144 0.0436 0.9999 0.6127 -2.250 -0.3809 0.06858 0.05977 0.0355 0.9999 0.5757 -2.000 -0.3452 0.06652 0.05768 0.0295 0.9999 0.5482 -1.750 -0.3062 0.06470 0.05579 0.0234 0.9999 0.5216 -1.500 -0.2644 0.06297 0.05404 0.0173 0.9999 0.4976 -1.250 -0.2221 0.06129 0.05240 0.0119 0.9999 0.4794 -0.750 -0.0023 0.05389 0.04455 -0.0240 0.7929 0.4651 -0.500 0.0322 0.05296 0.04335 -0.0270 0.6826 0.4662 -0.250 0.2154 0.04585 0.03393 -0.0509 0.6100 0.4949 0.000 0.3115 0.04422 0.03119 -0.0636 0.5830 0.5477 0.250 0.5148 0.04112 0.02808 -0.1001 0.5559 1.0001 0.750 0.6306 0.04429 0.02954 -0.1135 0.5400 1.0001 1.000 0.6710 0.04569 0.03054 -0.1165 0.5338 1.0001 1.250 0.7100 0.04720 0.03165 -0.1192 0.5277 1.0001 1.500 0.7489 0.04885 0.03292 -0.1218 0.5228 1.0001 1.750 0.7771 0.05040 0.03435 -0.1226 0.5204 1.0001 2.000 0.8022 0.05203 0.03596 -0.1230 0.5194 1.0001 2.250 0.8258 0.05381 0.03772 -0.1233 0.5197 1.0001 2.500 0.8475 0.05576 0.03969 -0.1233 0.5205 1.0001 2.750 0.8682 0.05794 0.04189 -0.1234 0.5217 1.0001 3.000 0.8719 0.06043 0.04461 -0.1212 0.5244 1.0001 3.250 0.8605 0.06374 0.04820 -0.1174 0.5287 1.0001 3.500 0.8429 0.06793 0.05260 -0.1137 0.5342 1.0001 3.750 0.8267 0.07254 0.05731 -0.1107 0.5403 1.0001 4.000 0.8291 0.07661 0.06134 -0.1100 0.5454 1.0001 4.250 0.7197 0.08726 0.07227 -0.1015 0.5645 1.0001 4.750 0.6451 0.10132 0.08634 -0.0998 0.5995 1.0001 5.000 0.6435 0.10621 0.09110 -0.1007 0.6150 1.0001