XFOIL Version 6.94 Calculated polar for: E193 (10.22%) 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.080 Re = 0.300 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- 0.000 0.4178 0.00822 0.00277 -0.0862 0.6843 1.0000 0.500 0.4679 0.00842 0.00281 -0.0850 0.6720 1.0000 1.000 0.5186 0.00866 0.00289 -0.0840 0.6605 1.0000 1.500 0.5696 0.00887 0.00300 -0.0830 0.6484 1.0000 2.000 0.6209 0.00911 0.00320 -0.0821 0.6365 1.0000 2.500 0.6728 0.00937 0.00341 -0.0812 0.6249 1.0000 3.000 0.7251 0.00963 0.00360 -0.0805 0.6122 1.0000 3.500 0.7766 0.00982 0.00382 -0.0795 0.5975 1.0000 4.000 0.8282 0.01002 0.00407 -0.0786 0.5819 1.0000 4.500 0.8796 0.01022 0.00429 -0.0776 0.5648 1.0000 5.000 0.9307 0.01042 0.00456 -0.0766 0.5465 1.0000 5.500 0.9808 0.01062 0.00485 -0.0754 0.5246 1.0000 6.000 1.0278 0.01078 0.00507 -0.0735 0.4856 1.0000 6.500 1.0727 0.01120 0.00543 -0.0714 0.4339 1.0000 7.000 1.1075 0.01234 0.00616 -0.0679 0.3301 1.0000 8.000 1.1506 0.01692 0.00945 -0.0582 0.0961 1.0000 8.500 1.1575 0.01982 0.01189 -0.0510 0.0267 1.0000 9.000 1.1722 0.02183 0.01403 -0.0450 0.0206 1.0000 9.500 1.1866 0.02404 0.01641 -0.0398 0.0187 1.0000 10.000 1.1924 0.02712 0.01965 -0.0348 0.0173 1.0000 10.500 1.1924 0.03123 0.02396 -0.0303 0.0164 1.0000 11.000 1.2046 0.03485 0.02780 -0.0273 0.0160 1.0000 11.500 1.2209 0.03858 0.03174 -0.0246 0.0157 1.0000 12.000 1.2417 0.04257 0.03601 -0.0220 0.0155 1.0000 12.500 1.2609 0.04720 0.04099 -0.0198 0.0154 1.0000 13.000 1.2686 0.05271 0.04692 -0.0180 0.0151 1.0000 13.500 1.2649 0.05935 0.05401 -0.0171 0.0149 1.0000 14.000 1.2479 0.06796 0.06311 -0.0172 0.0151 1.0000 14.500 1.2209 0.07822 0.07383 -0.0193 0.0153 1.0000 15.000 1.1886 0.09008 0.08610 -0.0235 0.0156 1.0000 15.500 1.1503 0.10440 0.10080 -0.0304 0.0158 1.0000 16.000 1.1117 0.12066 0.11738 -0.0397 0.0161 1.0000 16.500 1.0792 0.13759 0.13454 -0.0493 0.0167 1.0000