XFOIL Version 6.94 Calculated polar for: E224 (10.17%) 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 0.2340 0.00740 0.00282 -0.0542 0.8367 0.8812 0.500 0.2838 0.00738 0.00277 -0.0517 0.8080 0.9277 1.000 0.3507 0.00741 0.00270 -0.0532 0.7774 0.9652 1.500 0.4362 0.00747 0.00262 -0.0592 0.7424 0.9873 2.000 0.5044 0.00759 0.00258 -0.0619 0.7039 1.0000 2.500 0.5452 0.00779 0.00264 -0.0590 0.6635 1.0000 3.000 0.5895 0.00810 0.00278 -0.0566 0.6212 1.0000 3.500 0.6362 0.00847 0.00300 -0.0546 0.5744 1.0000 4.000 0.6840 0.00890 0.00328 -0.0529 0.5238 1.0000 4.500 0.7317 0.00942 0.00367 -0.0513 0.4677 1.0000 5.000 0.7786 0.01007 0.00411 -0.0497 0.4041 1.0000 5.500 0.8253 0.01082 0.00468 -0.0481 0.3427 1.0000 6.000 0.8697 0.01180 0.00539 -0.0463 0.2660 1.0000 6.500 0.9126 0.01301 0.00625 -0.0444 0.1864 1.0000 7.000 0.9553 0.01431 0.00729 -0.0426 0.1248 1.0000 7.500 0.9958 0.01582 0.00864 -0.0404 0.0747 1.0000 8.000 1.0308 0.01790 0.01061 -0.0372 0.0423 1.0000 8.500 1.0640 0.02000 0.01286 -0.0337 0.0334 1.0000 9.000 1.0887 0.02297 0.01595 -0.0294 0.0270 1.0000 9.500 1.1238 0.02471 0.01797 -0.0265 0.0236 1.0000 10.000 1.1460 0.02716 0.02055 -0.0222 0.0192 1.0000 10.500 1.1633 0.03120 0.02499 -0.0175 0.0171 1.0000 11.000 1.1776 0.03484 0.02901 -0.0129 0.0153 1.0000 11.500 1.1825 0.03909 0.03363 -0.0085 0.0142 1.0000 12.000 1.1801 0.04350 0.03825 -0.0053 0.0129 1.0000 12.500 1.1596 0.05117 0.04640 -0.0031 0.0125 1.0000 13.000 1.1226 0.06145 0.05719 -0.0040 0.0122 1.0000 13.500 1.0944 0.07171 0.06784 -0.0079 0.0123 1.0000 14.000 1.0534 0.08590 0.08242 -0.0159 0.0122 1.0000 14.500 1.0273 0.09967 0.09648 -0.0248 0.0123 1.0000 15.000 0.8851 0.15319 0.15053 -0.0565 0.0174 1.0000