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 = 0.045 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.4555 0.06258 0.05732 0.0403 0.9999 0.5520 -2.750 -0.2015 0.04397 0.03747 0.0069 0.4899 0.7145 -2.500 -0.2017 0.04332 0.03623 0.0083 0.4365 0.7045 -2.250 -0.1354 0.04521 0.03713 -0.0122 0.4069 0.5689 -2.000 -0.0109 0.04181 0.03079 -0.0436 0.3914 0.2360 -1.750 0.0147 0.04037 0.02902 -0.0431 0.3827 0.2326 -1.500 0.0461 0.03918 0.02732 -0.0436 0.3753 0.2325 -1.250 0.0837 0.03808 0.02570 -0.0453 0.3688 0.2360 -1.000 0.1262 0.03695 0.02416 -0.0480 0.3613 0.2387 -0.750 0.1781 0.03610 0.02284 -0.0527 0.3532 0.2439 -0.500 0.2398 0.03531 0.02153 -0.0592 0.3468 0.2548 -0.250 0.3199 0.03468 0.02063 -0.0700 0.3415 0.2812 0.000 0.4324 0.03323 0.01934 -0.0878 0.3370 0.4062 0.250 0.4957 0.03287 0.01938 -0.0953 0.3349 0.5090 0.500 0.5447 0.03259 0.01973 -0.0999 0.3339 0.6214 0.750 0.6594 0.03397 0.02129 -0.1191 0.3326 1.0001 1.000 0.6915 0.03489 0.02203 -0.1199 0.3329 1.0001 1.250 0.7215 0.03580 0.02284 -0.1202 0.3336 1.0001 1.500 0.7499 0.03671 0.02373 -0.1202 0.3347 1.0001 1.750 0.7771 0.03769 0.02474 -0.1200 0.3364 1.0001 2.000 0.8035 0.03879 0.02590 -0.1197 0.3379 1.0001 2.250 0.8285 0.03995 0.02712 -0.1192 0.3390 1.0001 2.500 0.8523 0.04120 0.02843 -0.1185 0.3400 1.0001 2.750 0.8751 0.04252 0.02983 -0.1177 0.3408 1.0001 3.000 0.8970 0.04393 0.03133 -0.1167 0.3420 1.0001 3.250 0.9178 0.04544 0.03298 -0.1157 0.3440 1.0001 3.500 0.9382 0.04713 0.03477 -0.1146 0.3467 1.0001 3.750 0.9593 0.04902 0.03672 -0.1138 0.3497 1.0001 4.000 0.9819 0.05130 0.03898 -0.1133 0.3522 1.0001 4.250 0.9895 0.05246 0.04079 -0.1105 0.3625 1.0001 4.500 1.0052 0.05484 0.04333 -0.1092 0.3686 1.0001 4.750 1.0276 0.05760 0.04608 -0.1090 0.3731 1.0001 5.000 1.0239 0.05987 0.04898 -0.1057 0.3897 1.0001