XFOIL Version 6.94 Calculated polar for: manu04 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.050 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -2.500 -0.0743 0.04703 0.03780 -0.0397 0.3775 0.3358 -2.250 -0.0800 0.04618 0.03724 -0.0309 0.3705 0.4220 -2.000 -0.0049 0.04054 0.02952 -0.0447 0.3640 0.2200 -1.750 0.0236 0.03888 0.02735 -0.0447 0.3574 0.2148 -1.500 0.0588 0.03747 0.02527 -0.0459 0.3509 0.2128 -1.250 0.0956 0.03653 0.02413 -0.0474 0.3445 0.2179 -1.000 0.1418 0.03551 0.02253 -0.0508 0.3373 0.2229 -0.750 0.2012 0.03466 0.02129 -0.0571 0.3309 0.2289 -0.500 0.2655 0.03387 0.02016 -0.0644 0.3274 0.2405 -0.250 0.3497 0.03337 0.01938 -0.0760 0.3238 0.2695 0.000 0.4468 0.03230 0.01858 -0.0908 0.3207 0.3857 0.250 0.5009 0.03222 0.01878 -0.0963 0.3194 0.4677 0.500 0.5454 0.03232 0.01923 -0.0999 0.3189 0.5273 0.750 0.6587 0.03290 0.02081 -0.1192 0.3182 1.0001 1.000 0.6911 0.03383 0.02157 -0.1200 0.3188 1.0001 1.250 0.7214 0.03480 0.02243 -0.1204 0.3198 1.0001 1.500 0.7503 0.03581 0.02337 -0.1205 0.3210 1.0001 1.750 0.7774 0.03686 0.02439 -0.1203 0.3218 1.0001 2.000 0.8033 0.03798 0.02552 -0.1199 0.3224 1.0001 2.250 0.8279 0.03913 0.02668 -0.1193 0.3228 1.0001 2.500 0.8517 0.04036 0.02794 -0.1186 0.3233 1.0001 2.750 0.8748 0.04167 0.02931 -0.1177 0.3242 1.0001 3.000 0.8972 0.04311 0.03083 -0.1168 0.3259 1.0001 3.250 0.9196 0.04471 0.03251 -0.1160 0.3284 1.0001 3.500 0.9422 0.04656 0.03436 -0.1154 0.3305 1.0001 3.750 0.9600 0.04763 0.03582 -0.1137 0.3364 1.0001 4.000 0.9753 0.04960 0.03811 -0.1120 0.3433 1.0001 4.250 0.9941 0.05177 0.04038 -0.1109 0.3483 1.0001 4.500 1.0168 0.05448 0.04307 -0.1107 0.3520 1.0001 4.750 1.0199 0.05631 0.04557 -0.1077 0.3683 1.0001