XFOIL Version 6.94 Calculated polar for: pp12 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.005 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.0261 0.06736 0.04575 -0.0225 0.9999 1.0001 -2.750 -0.0268 0.06616 0.04480 -0.0209 0.9999 1.0001 -2.500 -0.0280 0.06502 0.04395 -0.0192 0.9999 1.0001 -2.250 -0.0302 0.06387 0.04305 -0.0175 0.9999 1.0001 -2.000 -0.0326 0.06273 0.04220 -0.0156 0.9999 1.0001 -1.750 -0.0355 0.06157 0.04134 -0.0138 0.9999 1.0001 -1.500 -0.0377 0.06045 0.04046 -0.0120 0.9999 1.0001 -1.250 -0.0382 0.05941 0.03962 -0.0106 0.9999 1.0001 -1.000 -0.0351 0.05847 0.03882 -0.0100 0.9999 1.0001 -0.750 -0.0246 0.05775 0.03812 -0.0109 0.9999 1.0001 -0.500 -0.0038 0.05729 0.03756 -0.0140 0.9999 1.0001 -0.250 0.0264 0.05724 0.03735 -0.0188 0.9999 1.0001 0.000 0.0655 0.05758 0.03743 -0.0253 0.9999 1.0001 0.250 0.1146 0.05831 0.03782 -0.0337 0.9999 1.0001 0.500 0.1659 0.05936 0.03852 -0.0423 0.9999 1.0001 0.750 0.2241 0.06079 0.03953 -0.0521 0.9999 1.0001 1.000 0.2828 0.06244 0.04071 -0.0617 0.9999 1.0001 1.250 0.3394 0.06425 0.04210 -0.0705 0.9999 1.0001 1.500 0.3943 0.06619 0.04350 -0.0787 0.9999 1.0001 1.750 0.4465 0.06814 0.04490 -0.0857 0.9999 1.0001 2.000 0.4889 0.06995 0.04624 -0.0905 0.9999 1.0001 2.250 0.5301 0.07177 0.04746 -0.0944 0.9999 1.0001 2.500 0.5636 0.07343 0.04855 -0.0966 0.9999 1.0001 2.750 0.5923 0.07503 0.04966 -0.0977 0.9999 1.0001 3.000 0.6170 0.07659 0.05087 -0.0981 0.9999 1.0001 3.250 0.6402 0.07812 0.05216 -0.0982 0.9999 1.0001 3.500 0.6616 0.07970 0.05370 -0.0981 0.9999 1.0001 3.750 0.6820 0.08135 0.05544 -0.0980 0.9999 1.0001 4.000 0.7011 0.08311 0.05736 -0.0980 0.9999 1.0001 4.250 0.7199 0.08495 0.05957 -0.0979 0.9999 1.0001 4.500 0.7376 0.08694 0.06199 -0.0980 0.9999 1.0001 4.750 0.7545 0.08915 0.06471 -0.0982 0.9999 1.0001 5.000 0.7687 0.09167 0.06785 -0.0988 0.9999 1.0001