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 = 0.008 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.2973 0.06980 0.05407 0.0295 0.9999 0.8306 -2.750 -0.3009 0.06730 0.05172 0.0312 0.9999 0.8300 -2.500 -0.2856 0.06472 0.04922 0.0297 0.9999 0.8371 -2.250 -0.2619 0.06226 0.04683 0.0267 0.9999 0.8494 -2.000 -0.2203 0.05991 0.04447 0.0202 0.9999 0.8710 -1.750 -0.0723 0.05745 0.04187 -0.0089 0.9999 0.9795 -1.500 -0.0699 0.05580 0.04037 -0.0096 0.9999 1.0001 -1.250 -0.0751 0.05388 0.03857 -0.0080 0.9999 1.0001 -1.000 -0.0701 0.05233 0.03685 -0.0083 0.9999 1.0001 -0.750 -0.0477 0.05145 0.03553 -0.0117 0.9999 1.0001 -0.500 -0.0062 0.05131 0.03465 -0.0185 0.9999 1.0001 -0.250 0.0502 0.05186 0.03414 -0.0275 0.9999 1.0001 0.000 0.1065 0.05273 0.03387 -0.0351 0.9999 1.0001 0.250 0.1507 0.05353 0.03379 -0.0394 0.9999 1.0001 0.500 0.1857 0.05425 0.03401 -0.0416 0.9999 1.0001 0.750 0.2150 0.05507 0.03463 -0.0430 0.9999 1.0001 1.000 0.2378 0.05629 0.03588 -0.0442 0.9999 1.0001 1.250 0.2466 0.05862 0.03842 -0.0451 0.9999 1.0001 1.500 0.2342 0.06274 0.04266 -0.0457 0.9999 1.0001 1.750 0.2242 0.06681 0.04652 -0.0465 0.9999 1.0001 2.000 0.2231 0.07022 0.04967 -0.0475 0.9999 1.0001 2.250 0.2265 0.07327 0.05239 -0.0485 0.9999 1.0001 2.500 0.2323 0.07616 0.05496 -0.0495 0.9999 1.0001 2.750 0.2396 0.07891 0.05740 -0.0504 0.9999 1.0001 3.000 0.2480 0.08159 0.05976 -0.0512 0.9999 1.0001 3.250 0.2573 0.08419 0.06206 -0.0520 0.9999 1.0001 3.500 0.2671 0.08678 0.06436 -0.0528 0.9999 1.0001 3.750 0.2773 0.08933 0.06663 -0.0536 0.9999 1.0001 4.000 0.2877 0.09189 0.06892 -0.0543 0.9999 1.0001 4.250 0.2985 0.09443 0.07120 -0.0550 0.9999 1.0001 4.500 0.3093 0.09699 0.07352 -0.0556 0.9999 1.0001 4.750 0.3203 0.09952 0.07582 -0.0563 0.9999 1.0001 5.000 0.3314 0.10208 0.07816 -0.0569 0.9999 1.0001