XFOIL Version 6.94 Calculated polar for: E174 (Dicke 8.92%) 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 0.4029 0.00669 0.00193 -0.0818 0.6543 1.0000 0.500 0.4585 0.00685 0.00192 -0.0817 0.6429 1.0000 1.000 0.5141 0.00705 0.00195 -0.0815 0.6319 1.0000 1.500 0.5699 0.00719 0.00202 -0.0815 0.6201 1.0000 2.000 0.6256 0.00737 0.00214 -0.0814 0.6085 1.0000 2.500 0.6810 0.00757 0.00228 -0.0813 0.5959 1.0000 3.000 0.7362 0.00776 0.00241 -0.0811 0.5815 1.0000 3.500 0.7911 0.00794 0.00259 -0.0809 0.5660 1.0000 4.000 0.8459 0.00813 0.00279 -0.0807 0.5496 1.0000 4.500 0.9003 0.00828 0.00302 -0.0804 0.5278 1.0000 5.000 0.9535 0.00851 0.00323 -0.0799 0.4957 1.0000 5.500 1.0052 0.00890 0.00356 -0.0792 0.4493 1.0000 6.000 1.0510 0.00991 0.00414 -0.0778 0.3437 1.0000 6.500 1.0887 0.01193 0.00543 -0.0756 0.1981 1.0000 7.000 1.1217 0.01438 0.00710 -0.0729 0.0648 1.0000 7.500 1.1571 0.01643 0.00890 -0.0701 0.0184 1.0000 8.000 1.1965 0.01787 0.01050 -0.0678 0.0155 1.0000 8.500 1.2249 0.02006 0.01292 -0.0640 0.0140 1.0000 9.000 1.2490 0.02217 0.01523 -0.0597 0.0135 1.0000 9.500 1.2621 0.02453 0.01779 -0.0540 0.0131 1.0000 10.000 1.2739 0.02748 0.02097 -0.0492 0.0129 1.0000 10.500 1.2884 0.03093 0.02465 -0.0453 0.0126 1.0000 11.000 1.3078 0.03492 0.02892 -0.0421 0.0126 1.0000 11.500 1.3287 0.03960 0.03396 -0.0393 0.0127 1.0000 12.000 1.3390 0.04476 0.03952 -0.0367 0.0125 1.0000 12.500 1.3360 0.05147 0.04672 -0.0345 0.0126 1.0000 13.000 1.3184 0.05999 0.05572 -0.0335 0.0130 1.0000 13.500 1.2899 0.07009 0.06626 -0.0345 0.0133 1.0000 15.000 0.9792 0.10434 0.10197 -0.0504 0.0134 1.0000 15.500 0.9493 0.11786 0.11572 -0.0586 0.0141 1.0000 16.000 0.9170 0.13367 0.13177 -0.0697 0.0143 1.0000