XFOIL Version 6.94 Calculated polar for: RG-15 8.9% 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.080 Re = 0.450 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- 0.000 0.2669 0.00594 0.00186 -0.0607 0.8471 0.8426 0.500 0.3184 0.00597 0.00181 -0.0591 0.8093 0.8715 1.000 0.3668 0.00604 0.00181 -0.0568 0.7687 0.9034 1.500 0.4137 0.00614 0.00182 -0.0542 0.7253 0.9414 2.000 0.4785 0.00632 0.00185 -0.0558 0.6752 0.9770 2.500 0.5485 0.00661 0.00194 -0.0591 0.6178 1.0000 3.000 0.5952 0.00700 0.00210 -0.0574 0.5623 1.0000 3.500 0.6443 0.00748 0.00234 -0.0562 0.5010 1.0000 4.000 0.6939 0.00801 0.00268 -0.0551 0.4383 1.0000 4.500 0.7433 0.00864 0.00307 -0.0540 0.3714 1.0000 5.000 0.7918 0.00941 0.00359 -0.0528 0.2963 1.0000 5.500 0.8398 0.01029 0.00419 -0.0517 0.2236 1.0000 6.000 0.8859 0.01140 0.00495 -0.0503 0.1456 1.0000 6.500 0.9318 0.01257 0.00588 -0.0489 0.0877 1.0000 7.000 0.9773 0.01376 0.00690 -0.0475 0.0501 1.0000 7.500 1.0198 0.01532 0.00834 -0.0454 0.0190 1.0000 8.000 1.0597 0.01718 0.01036 -0.0427 0.0121 1.0000 9.000 1.1219 0.02267 0.01647 -0.0350 0.0097 1.0000 9.500 1.1551 0.02501 0.01908 -0.0319 0.0086 1.0000 10.000 1.1808 0.02828 0.02271 -0.0280 0.0081 1.0000 10.500 1.1964 0.03213 0.02697 -0.0230 0.0077 1.0000 11.000 1.1975 0.03632 0.03158 -0.0168 0.0074 1.0000 11.500 1.1875 0.04123 0.03691 -0.0114 0.0072 1.0000 12.000 1.1657 0.04785 0.04398 -0.0082 0.0071 1.0000 12.500 1.1333 0.05711 0.05369 -0.0088 0.0071 1.0000 13.000 1.0923 0.07006 0.06708 -0.0147 0.0073 1.0000 13.500 1.0415 0.08847 0.08592 -0.0265 0.0077 1.0000 14.000 0.9903 0.11133 0.10909 -0.0417 0.0080 1.0000 14.500 0.9025 0.15083 0.14870 -0.0624 0.0096 1.0000 15.000 0.8889 0.16658 0.16441 -0.0699 0.0106 1.0000