XFOIL Version 6.94 Calculated polar for: manu4/38 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.070 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.0679 0.04081 0.03063 -0.0543 0.3140 0.1768 -2.750 -0.0495 0.03926 0.02890 -0.0526 0.3091 0.1754 -2.500 -0.0278 0.03760 0.02697 -0.0513 0.3050 0.1746 -2.250 -0.0023 0.03597 0.02501 -0.0507 0.3011 0.1736 -2.000 0.0295 0.03431 0.02291 -0.0513 0.2972 0.1717 -1.750 0.0769 0.03269 0.02069 -0.0550 0.2932 0.1709 -1.500 0.1290 0.03135 0.01887 -0.0597 0.2900 0.1722 -1.250 0.1815 0.03039 0.01744 -0.0644 0.2875 0.1771 -1.000 0.2296 0.03018 0.01717 -0.0685 0.2854 0.1849 -0.750 0.2805 0.02977 0.01646 -0.0731 0.2833 0.1932 -0.500 0.3263 0.02966 0.01621 -0.0766 0.2819 0.2051 -0.250 0.3876 0.02965 0.01621 -0.0839 0.2806 0.2308 0.000 0.4446 0.02974 0.01644 -0.0903 0.2793 0.3013 0.250 0.4907 0.03000 0.01680 -0.0944 0.2776 0.3710 0.500 0.5300 0.03051 0.01740 -0.0970 0.2757 0.4157 0.750 0.5681 0.03131 0.01824 -0.0995 0.2740 0.4489 1.000 0.6020 0.03171 0.01886 -0.1008 0.2735 0.4828 1.250 0.6360 0.03200 0.01966 -0.1024 0.2737 0.5559 1.500 0.7523 0.03362 0.02191 -0.1225 0.2742 1.0001 1.750 0.7791 0.03429 0.02256 -0.1221 0.2752 1.0001 2.000 0.8043 0.03505 0.02343 -0.1212 0.2770 1.0001 2.250 0.8290 0.03609 0.02462 -0.1204 0.2794 1.0001 2.500 0.8533 0.03730 0.02595 -0.1196 0.2823 1.0001 2.750 0.8772 0.03866 0.02738 -0.1188 0.2849 1.0001 3.000 0.9003 0.04012 0.02891 -0.1180 0.2876 1.0001 3.250 0.9234 0.04174 0.03053 -0.1173 0.2900 1.0001 3.500 0.9477 0.04376 0.03251 -0.1169 0.2920 1.0001 3.750 0.9630 0.04491 0.03435 -0.1143 0.3043 1.0001 4.000 0.9844 0.04709 0.03655 -0.1136 0.3088 1.0001 4.250 0.9998 0.04927 0.03925 -0.1118 0.3256 1.0001 4.500 1.0243 0.05177 0.04170 -0.1117 0.3314 1.0001