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 -0.2863 0.06419 0.05830 0.0437 0.9999 0.8660 -2.750 -0.3571 0.06192 0.05615 0.0574 0.9999 0.8447 -2.500 -0.4121 0.05905 0.05339 0.0673 0.9999 0.8219 -2.250 -0.1943 0.05590 0.04912 -0.0068 0.8755 0.4870 -2.000 -0.1363 0.05268 0.04537 -0.0168 0.7609 0.3910 -1.750 -0.0121 0.04837 0.03808 -0.0358 0.4861 0.3301 -1.500 0.0220 0.04719 0.03619 -0.0378 0.4694 0.3220 -1.250 0.0604 0.04609 0.03444 -0.0402 0.4567 0.3199 -1.000 0.1042 0.04489 0.03266 -0.0434 0.4467 0.3197 -0.750 0.1626 0.04375 0.03082 -0.0496 0.4381 0.3204 -0.500 0.2246 0.04272 0.02926 -0.0564 0.4303 0.3275 -0.250 0.2977 0.04202 0.02787 -0.0655 0.4215 0.3462 0.000 0.3742 0.04125 0.02679 -0.0754 0.4130 0.3893 0.250 0.4713 0.03929 0.02521 -0.0893 0.4045 0.5367 0.500 0.6146 0.03880 0.02564 -0.1142 0.3988 1.0001 0.750 0.6662 0.03993 0.02636 -0.1193 0.3978 1.0001 1.000 0.7091 0.04101 0.02718 -0.1224 0.3975 1.0001 1.250 0.7462 0.04210 0.02810 -0.1244 0.3978 1.0001 1.500 0.7784 0.04317 0.02910 -0.1254 0.3987 1.0001 1.750 0.8074 0.04431 0.03024 -0.1258 0.4002 1.0001 2.000 0.8342 0.04554 0.03153 -0.1259 0.4021 1.0001 2.250 0.8591 0.04690 0.03295 -0.1257 0.4045 1.0001 2.500 0.8828 0.04838 0.03451 -0.1253 0.4072 1.0001 2.750 0.9056 0.05001 0.03619 -0.1249 0.4102 1.0001 3.000 0.9269 0.05179 0.03804 -0.1243 0.4133 1.0001 3.250 0.9482 0.05372 0.04003 -0.1237 0.4160 1.0001 3.500 0.9697 0.05586 0.04216 -0.1232 0.4182 1.0001 3.750 0.9769 0.05737 0.04402 -0.1206 0.4228 1.0001 4.000 0.9800 0.05964 0.04655 -0.1177 0.4277 1.0001 4.250 0.9839 0.06227 0.04936 -0.1151 0.4323 1.0001 4.500 0.9907 0.06511 0.05230 -0.1131 0.4367 1.0001 4.750 1.0046 0.06807 0.05531 -0.1122 0.4406 1.0001 5.000 0.9523 0.07282 0.06049 -0.1045 0.4542 1.0001