XFOIL Version 6.94 Calculated polar for: manu4/38 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.040 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.4945 0.06103 0.05562 0.0614 0.9999 0.6826 -2.750 -0.4905 0.05778 0.05246 0.0644 0.9999 0.7055 -2.500 -0.4817 0.05459 0.04936 0.0659 0.9999 0.7208 -2.250 -0.4691 0.05198 0.04679 0.0651 0.9999 0.7265 -1.750 0.0081 0.04143 0.03026 -0.0420 0.4216 0.2535 -1.500 0.0357 0.04046 0.02873 -0.0419 0.4079 0.2548 -1.250 0.0696 0.03943 0.02713 -0.0430 0.3964 0.2559 -1.000 0.1098 0.03831 0.02549 -0.0452 0.3868 0.2574 -0.750 0.1553 0.03732 0.02418 -0.0486 0.3778 0.2630 -0.500 0.2259 0.03671 0.02273 -0.0569 0.3710 0.2764 -0.250 0.2971 0.03601 0.02181 -0.0657 0.3660 0.3091 0.000 0.4121 0.03403 0.02010 -0.0838 0.3603 0.4532 0.250 0.4780 0.03305 0.01995 -0.0916 0.3575 0.6010 0.500 0.6102 0.03379 0.02094 -0.1145 0.3544 1.0001 0.750 0.6503 0.03469 0.02155 -0.1169 0.3539 1.0001 1.000 0.6857 0.03559 0.02227 -0.1184 0.3540 1.0001 1.250 0.7180 0.03653 0.02312 -0.1192 0.3541 1.0001 1.500 0.7479 0.03755 0.02406 -0.1196 0.3539 1.0001 1.750 0.7755 0.03866 0.02510 -0.1196 0.3533 1.0001 2.000 0.8021 0.03983 0.02624 -0.1194 0.3529 1.0001 2.250 0.8276 0.04106 0.02746 -0.1191 0.3524 1.0001 2.500 0.8519 0.04222 0.02872 -0.1185 0.3533 1.0001 2.750 0.8746 0.04345 0.03012 -0.1176 0.3553 1.0001 3.000 0.8964 0.04486 0.03170 -0.1167 0.3583 1.0001 3.250 0.9172 0.04642 0.03341 -0.1158 0.3617 1.0001 3.500 0.9373 0.04814 0.03529 -0.1148 0.3651 1.0001 3.750 0.9579 0.05004 0.03726 -0.1139 0.3682 1.0001 4.000 0.9799 0.05224 0.03946 -0.1134 0.3709 1.0001 4.250 0.9871 0.05357 0.04137 -0.1107 0.3798 1.0001 4.500 1.0006 0.05590 0.04390 -0.1092 0.3863 1.0001 4.750 1.0206 0.05851 0.04653 -0.1087 0.3912 1.0001 5.000 1.0144 0.06101 0.04957 -0.1053 0.4050 1.0001