XFOIL Version 6.94 Calculated polar for: manu4/36 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 -0.4392 0.06964 0.06033 0.0410 0.9999 0.6073 -2.750 -0.4113 0.06759 0.05807 0.0339 0.9999 0.5718 -2.500 -0.3812 0.06546 0.05575 0.0279 0.9999 0.5443 -2.250 -0.3480 0.06360 0.05366 0.0221 0.9999 0.5193 -2.000 -0.3127 0.06178 0.05162 0.0168 0.9999 0.4973 -1.750 -0.2774 0.05989 0.04954 0.0127 0.9999 0.4799 -1.500 -0.2406 0.05819 0.04764 0.0087 0.9999 0.4663 -1.250 -0.2015 0.05659 0.04589 0.0047 0.9999 0.4564 -1.000 -0.1620 0.05503 0.04428 0.0013 0.9999 0.4551 -0.750 -0.1214 0.05339 0.04275 -0.0019 0.9999 0.4590 -0.500 -0.0790 0.05221 0.04170 -0.0061 0.9999 0.4634 -0.250 0.1007 0.04684 0.03600 -0.0363 0.7987 0.4908 0.000 0.1633 0.04468 0.03371 -0.0426 0.6818 0.5279 0.250 0.4804 0.03787 0.02424 -0.0945 0.5879 1.0001 0.500 0.5372 0.03921 0.02450 -0.1004 0.5736 1.0001 0.750 0.5837 0.04042 0.02507 -0.1043 0.5641 1.0001 1.000 0.6302 0.04174 0.02579 -0.1083 0.5570 1.0001 1.250 0.6658 0.04293 0.02680 -0.1103 0.5520 1.0001 1.500 0.6998 0.04424 0.02793 -0.1121 0.5472 1.0001 1.750 0.7330 0.04570 0.02922 -0.1137 0.5421 1.0001 2.000 0.7658 0.04725 0.03061 -0.1153 0.5377 1.0001 2.250 0.7950 0.04892 0.03218 -0.1163 0.5344 1.0001 2.500 0.8156 0.05064 0.03400 -0.1160 0.5321 1.0001 2.750 0.8348 0.05252 0.03596 -0.1156 0.5311 1.0001 3.000 0.8521 0.05462 0.03815 -0.1151 0.5314 1.0001 3.250 0.8686 0.05694 0.04060 -0.1148 0.5328 1.0001 3.500 0.8830 0.05953 0.04327 -0.1143 0.5347 1.0001 3.750 0.8701 0.06308 0.04714 -0.1109 0.5390 1.0001 4.000 0.8416 0.06799 0.05229 -0.1067 0.5457 1.0001 4.250 0.8235 0.07295 0.05732 -0.1042 0.5524 1.0001 4.750 0.7359 0.08676 0.07115 -0.0979 0.5761 1.0001