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 = 0.030 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 0.2983 0.06925 0.05961 -0.0605 0.3827 0.8811 -2.750 0.2414 0.06804 0.05863 -0.0490 0.3838 0.8453 -2.500 0.1782 0.06647 0.05730 -0.0368 0.3853 0.8150 -2.250 0.0998 0.06470 0.05581 -0.0210 0.3871 0.7967 -2.000 0.0254 0.06257 0.05391 -0.0067 0.3891 0.7772 -1.750 0.0131 0.06113 0.05238 -0.0089 0.3894 0.7112 -1.500 0.1492 0.06413 0.05423 -0.0482 0.3870 0.4910 -1.250 0.2437 0.06353 0.05277 -0.0673 0.3860 0.3850 -1.000 0.3046 0.06218 0.05102 -0.0756 0.3859 0.3548 -0.750 0.3589 0.06125 0.04973 -0.0820 0.3864 0.3352 -0.500 0.4075 0.06072 0.04880 -0.0868 0.3875 0.3202 -0.250 0.4488 0.06004 0.04799 -0.0899 0.3890 0.3163 0.000 0.4944 0.05975 0.04749 -0.0939 0.3901 0.3149 0.250 0.5393 0.05995 0.04739 -0.0977 0.3906 0.3179 0.500 0.5842 0.06011 0.04748 -0.1019 0.3912 0.3276 0.750 0.6272 0.06077 0.04793 -0.1055 0.3918 0.3397 1.000 0.6764 0.06145 0.04857 -0.1107 0.3923 0.3559 1.250 0.7307 0.06241 0.04953 -0.1172 0.3953 0.3851 1.500 0.8064 0.06304 0.05076 -0.1282 0.4006 0.4940 1.750 0.9280 0.06553 0.05442 -0.1497 0.4068 1.0001 2.000 0.9224 0.06659 0.05577 -0.1444 0.4141 1.0001 2.250 0.9131 0.06917 0.05854 -0.1392 0.4227 1.0001 2.500 0.9186 0.07231 0.06160 -0.1368 0.4298 1.0001 2.750 0.9569 0.07604 0.06495 -0.1400 0.4361 1.0001 3.000 0.8265 0.08051 0.07052 -0.1178 0.4528 1.0001 3.250 0.7559 0.08688 0.07732 -0.1078 0.4700 1.0001 3.500 0.7393 0.09428 0.08475 -0.1074 0.4933 1.0001 4.250 0.0652 0.12552 0.11610 -0.0401 0.8701 0.3022 4.500 0.0962 0.12906 0.11934 -0.0448 0.8735 0.3041 4.750 0.1182 0.13030 0.12034 -0.0476 0.8672 0.3075 5.000 0.1588 0.13442 0.12410 -0.0537 0.8597 0.3159