XFOIL Version 6.94 Calculated polar for: Profil 387 Dicke 9.06% 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 0.3983 0.00674 0.00192 -0.0804 0.6494 1.0000 0.500 0.4536 0.00687 0.00192 -0.0803 0.6377 1.0000 1.000 0.5089 0.00706 0.00196 -0.0801 0.6267 1.0000 1.500 0.5642 0.00725 0.00203 -0.0799 0.6156 1.0000 2.000 0.6197 0.00740 0.00216 -0.0798 0.6038 1.0000 2.500 0.6749 0.00759 0.00229 -0.0797 0.5911 1.0000 3.000 0.7299 0.00778 0.00243 -0.0795 0.5773 1.0000 3.500 0.7848 0.00797 0.00261 -0.0792 0.5623 1.0000 4.000 0.8393 0.00815 0.00281 -0.0790 0.5461 1.0000 4.500 0.8936 0.00833 0.00305 -0.0786 0.5261 1.0000 5.000 0.9465 0.00854 0.00324 -0.0780 0.4944 1.0000 5.500 0.9983 0.00889 0.00356 -0.0774 0.4530 1.0000 6.000 1.0448 0.00976 0.00407 -0.0760 0.3589 1.0000 6.500 1.0828 0.01166 0.00527 -0.0738 0.2164 1.0000 7.000 1.1152 0.01410 0.00692 -0.0710 0.0791 1.0000 7.500 1.1479 0.01635 0.00879 -0.0678 0.0186 1.0000 8.000 1.1877 0.01770 0.01031 -0.0656 0.0155 1.0000 9.000 1.2383 0.02196 0.01497 -0.0571 0.0132 1.0000 9.500 1.2510 0.02428 0.01751 -0.0514 0.0130 1.0000 10.000 1.2621 0.02719 0.02063 -0.0466 0.0127 1.0000 10.500 1.2752 0.03061 0.02428 -0.0427 0.0124 1.0000 11.000 1.2931 0.03451 0.02843 -0.0394 0.0124 1.0000 11.500 1.3138 0.03887 0.03311 -0.0367 0.0124 1.0000 12.000 1.3270 0.04379 0.03843 -0.0342 0.0122 1.0000 12.500 1.3286 0.04999 0.04506 -0.0321 0.0123 1.0000 13.000 1.3162 0.05784 0.05338 -0.0309 0.0126 1.0000 13.500 1.2900 0.06775 0.06375 -0.0315 0.0129 1.0000 15.000 0.9946 0.09851 0.09594 -0.0433 0.0137 1.0000 15.500 0.9653 0.11100 0.10871 -0.0518 0.0138 1.0000 16.000 0.9290 0.12676 0.12476 -0.0629 0.0140 1.0000