XFOIL Version 6.94 Calculated polar for: manu4/25 2 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 -0.1860 0.10247 0.09130 0.0285 0.9999 0.8584 -2.750 -0.1960 0.09965 0.08881 0.0311 0.9999 0.8477 -2.500 -0.2179 0.09687 0.08632 0.0356 0.9999 0.8370 -2.250 -0.2428 0.09396 0.08371 0.0407 0.9999 0.8264 -2.000 -0.2711 0.09103 0.08105 0.0462 0.9999 0.8169 -1.750 -0.2883 0.08823 0.07856 0.0495 0.9999 0.8085 -1.500 -0.3122 0.08559 0.07612 0.0536 0.9999 0.7998 -1.250 -0.3113 0.08389 0.07477 0.0531 0.9999 0.7955 -1.000 -0.3171 0.08341 0.07465 0.0528 0.9999 0.7914 -0.750 -0.3363 0.08491 0.07636 0.0533 0.9999 0.7869 -0.500 -0.3539 0.08696 0.07837 0.0532 0.9999 0.7833 -0.250 -0.3596 0.08858 0.07986 0.0516 0.9999 0.7808 0.000 -0.3555 0.08993 0.08103 0.0488 0.9999 0.7798 0.250 -0.3450 0.09122 0.08216 0.0451 0.9999 0.7806 0.500 -0.3295 0.09252 0.08327 0.0409 0.9999 0.7825 0.750 -0.3103 0.09387 0.08446 0.0362 0.9999 0.7859 1.000 -0.2877 0.09532 0.08576 0.0310 0.9999 0.7917 1.500 -0.2317 0.09847 0.08893 0.0189 0.9999 0.8146 1.750 -0.1936 0.10025 0.09102 0.0107 0.9999 0.8385 2.000 -0.1526 0.10140 0.09245 0.0007 0.9999 0.8949 2.250 -0.1241 0.10440 0.09488 -0.0058 0.9999 1.0001 2.500 -0.1003 0.10662 0.09657 -0.0109 0.9999 1.0001 2.750 -0.0734 0.10929 0.09867 -0.0166 0.9999 1.0001 3.000 -0.0438 0.11230 0.10108 -0.0226 0.9999 1.0001 3.250 -0.0121 0.11563 0.10374 -0.0288 0.9999 1.0001 3.500 0.0206 0.11920 0.10663 -0.0351 0.9999 1.0001 3.750 0.0542 0.12298 0.10969 -0.0414 0.9999 1.0001 4.000 0.0892 0.12699 0.11291 -0.0477 0.9999 1.0001 4.250 0.1239 0.13112 0.11616 -0.0537 0.9999 1.0001 4.500 0.1558 0.13514 0.11922 -0.0587 0.9999 1.0001 4.750 0.1826 0.13882 0.12196 -0.0624 0.9999 1.0001 5.000 0.2046 0.14216 0.12447 -0.0649 0.9999 1.0001