XFOIL Version 6.94 Calculated polar for: manu4/28 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 -0.1319 0.07158 0.06315 -0.0132 0.3567 0.4924 -2.750 -0.1089 0.06791 0.05939 -0.0113 0.3518 0.5441 -2.500 -0.0938 0.06436 0.05589 -0.0080 0.3494 0.5938 -2.250 -0.0756 0.06082 0.05237 -0.0062 0.3473 0.6370 -2.000 -0.0554 0.05791 0.04939 -0.0062 0.3461 0.6688 -1.750 -0.0159 0.05441 0.04583 -0.0100 0.3445 0.6905 -1.500 0.0343 0.05203 0.04331 -0.0172 0.3428 0.6955 -1.250 0.1246 0.05220 0.04305 -0.0359 0.3411 0.6418 -1.000 0.3317 0.05353 0.04168 -0.0851 0.3391 0.2577 -0.750 0.3756 0.05223 0.04025 -0.0886 0.3392 0.2549 -0.500 0.4189 0.05153 0.03923 -0.0917 0.3397 0.2571 -0.250 0.4578 0.05096 0.03864 -0.0941 0.3407 0.2625 0.000 0.4972 0.05072 0.03828 -0.0965 0.3419 0.2677 0.250 0.5427 0.05077 0.03802 -0.1001 0.3436 0.2749 0.500 0.5954 0.05109 0.03821 -0.1056 0.3459 0.2882 0.750 0.6547 0.05171 0.03851 -0.1124 0.3480 0.3180 1.000 0.7329 0.05193 0.03898 -0.1237 0.3495 0.4159 1.500 0.8109 0.05430 0.04187 -0.1295 0.3517 0.5355 1.750 0.8404 0.05582 0.04374 -0.1304 0.3527 0.5913 2.000 0.8646 0.05561 0.04417 -0.1300 0.3569 0.6532 2.250 0.9392 0.05817 0.04796 -0.1423 0.3701 1.0001 2.500 0.9618 0.06088 0.05059 -0.1423 0.3761 1.0001 2.750 0.9957 0.06470 0.05413 -0.1446 0.3812 1.0001 3.000 0.9756 0.06530 0.05545 -0.1372 0.3992 1.0001 3.250 1.0099 0.06941 0.05926 -0.1397 0.4076 1.0001 3.500 0.9748 0.07184 0.06233 -0.1317 0.4312 1.0001 3.750 0.9362 0.07681 0.06771 -0.1252 0.4598 1.0001 4.000 0.1533 0.11181 0.10282 -0.0517 0.8475 0.2621 4.250 0.1778 0.11390 0.10472 -0.0547 0.8372 0.2682 4.500 0.2342 0.11859 0.10906 -0.0637 0.8254 0.2822 4.750 0.2667 0.12069 0.11094 -0.0681 0.8075 0.2994 5.000 0.3190 0.12437 0.11459 -0.0774 0.7894 0.3437