XFOIL Version 6.94 Calculated polar for: manu01 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.030 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.2925 0.06452 0.05859 0.0449 0.9999 0.8670 -2.750 -0.3436 0.06193 0.05614 0.0546 0.9999 0.8428 -2.500 -0.4113 0.05922 0.05353 0.0673 0.9999 0.8231 -2.250 -0.4601 0.05626 0.05062 0.0741 0.9999 0.7947 -2.000 -0.4181 0.05598 0.05019 0.0553 0.9981 0.6883 -1.500 0.0227 0.04738 0.03631 -0.0382 0.4719 0.3216 -1.250 0.0577 0.04602 0.03441 -0.0398 0.4552 0.3173 -1.000 0.1013 0.04486 0.03260 -0.0431 0.4416 0.3145 -0.750 0.1555 0.04379 0.03088 -0.0483 0.4317 0.3170 -0.500 0.2168 0.04290 0.02940 -0.0550 0.4238 0.3313 -0.250 0.2938 0.04217 0.02796 -0.0649 0.4169 0.3529 0.000 0.3768 0.04109 0.02659 -0.0760 0.4118 0.3944 0.250 0.4736 0.03877 0.02497 -0.0898 0.4075 0.5732 0.500 0.6156 0.03869 0.02552 -0.1145 0.4033 1.0001 0.750 0.6650 0.03973 0.02622 -0.1191 0.4027 1.0001 1.000 0.7071 0.04077 0.02702 -0.1221 0.4028 1.0001 1.250 0.7439 0.04185 0.02795 -0.1241 0.4030 1.0001 1.500 0.7767 0.04299 0.02899 -0.1252 0.4029 1.0001 1.750 0.8064 0.04422 0.03015 -0.1258 0.4026 1.0001 2.000 0.8335 0.04552 0.03143 -0.1259 0.4022 1.0001 2.250 0.8586 0.04689 0.03281 -0.1256 0.4023 1.0001 2.500 0.8818 0.04833 0.03430 -0.1251 0.4027 1.0001 2.750 0.9032 0.04988 0.03593 -0.1244 0.4041 1.0001 3.000 0.9232 0.05157 0.03775 -0.1235 0.4071 1.0001 3.250 0.9419 0.05343 0.03971 -0.1225 0.4104 1.0001 3.500 0.9611 0.05547 0.04180 -0.1217 0.4137 1.0001 3.750 0.9826 0.05774 0.04406 -0.1213 0.4168 1.0001 4.000 0.9842 0.05951 0.04624 -0.1181 0.4232 1.0001 4.250 0.9843 0.06217 0.04916 -0.1151 0.4304 1.0001 4.500 0.9946 0.06501 0.05208 -0.1137 0.4363 1.0001 4.750 0.9875 0.06811 0.05543 -0.1104 0.4448 1.0001 5.000 0.9603 0.07274 0.06030 -0.1057 0.4571 1.0001