XFOIL Version 6.94 Calculated polar for: pp13 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.040 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.2139 0.06931 0.06208 -0.0192 0.9998 0.2901 -2.250 -0.1682 0.06288 0.05577 -0.0189 0.9998 0.3541 -2.000 -0.1570 0.06050 0.05347 -0.0170 0.9998 0.3816 -1.750 -0.1455 0.05834 0.05139 -0.0152 0.9998 0.4205 -1.500 -0.1392 0.05589 0.04905 -0.0114 0.9998 0.4673 -0.750 -0.1369 0.04779 0.04135 0.0060 0.9998 0.6523 -0.500 -0.1315 0.04503 0.03868 0.0106 0.9998 0.6970 -0.250 -0.1184 0.04278 0.03649 0.0132 0.9998 0.7383 0.000 -0.0971 0.04072 0.03447 0.0137 0.9998 0.7724 0.250 -0.0620 0.03936 0.03308 0.0106 0.9998 0.8023 0.500 -0.0159 0.03870 0.03240 0.0045 0.9998 0.8213 0.750 0.0430 0.03891 0.03256 -0.0053 0.9998 0.8190 1.000 0.1068 0.04060 0.03415 -0.0176 0.9998 0.7729 1.250 0.1695 0.04392 0.03723 -0.0317 0.9998 0.6403 1.500 0.2281 0.04642 0.03934 -0.0435 0.9998 0.4914 1.750 0.2807 0.04782 0.04025 -0.0516 0.9998 0.3675 2.000 0.7240 0.03496 0.02245 -0.0999 0.1944 0.2351 2.250 0.7622 0.03594 0.02309 -0.1019 0.1824 0.2359 2.500 0.8023 0.03687 0.02379 -0.1041 0.1752 0.2435 2.750 0.8576 0.03792 0.02465 -0.1089 0.1685 0.2596 3.000 0.9384 0.03840 0.02536 -0.1176 0.1605 0.2970 3.250 1.0437 0.03901 0.02657 -0.1299 0.1567 1.0002 3.500 1.0847 0.04052 0.02781 -0.1304 0.1571 1.0002 3.750 1.1210 0.04232 0.02965 -0.1305 0.1587 1.0002 4.000 1.1538 0.04446 0.03195 -0.1302 0.1610 1.0002 4.250 1.1842 0.04688 0.03453 -0.1297 0.1640 1.0002 4.500 1.2129 0.04979 0.03757 -0.1291 0.1669 1.0002 4.750 1.2394 0.05312 0.04104 -0.1285 0.1687 1.0002 5.000 1.2636 0.05605 0.04426 -0.1274 0.1699 1.0002