XFOIL Version 6.94 Calculated polar for: AG36 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 0.3992 0.00664 0.00151 -0.0451 0.5916 1.0000 0.500 0.4537 0.00697 0.00166 -0.0448 0.5625 1.0000 1.000 0.5087 0.00726 0.00183 -0.0446 0.5363 1.0000 1.500 0.5638 0.00753 0.00203 -0.0445 0.5097 1.0000 2.000 0.6189 0.00779 0.00223 -0.0444 0.4806 1.0000 2.500 0.6739 0.00806 0.00246 -0.0443 0.4433 1.0000 3.000 0.7281 0.00846 0.00273 -0.0442 0.3959 1.0000 3.500 0.7812 0.00904 0.00314 -0.0440 0.3400 1.0000 4.000 0.8337 0.00973 0.00366 -0.0437 0.2829 1.0000 4.500 0.8850 0.01060 0.00428 -0.0434 0.2171 1.0000 5.000 0.9356 0.01157 0.00504 -0.0431 0.1604 1.0000 5.500 0.9854 0.01261 0.00590 -0.0427 0.1119 1.0000 6.000 1.0333 0.01391 0.00699 -0.0421 0.0625 1.0000 6.500 1.0774 0.01577 0.00867 -0.0408 0.0278 1.0000 7.000 1.1200 0.01773 0.01077 -0.0392 0.0193 1.0000 7.500 1.1590 0.01992 0.01311 -0.0373 0.0154 1.0000 8.000 1.1976 0.02197 0.01538 -0.0353 0.0134 1.0000 8.500 1.2240 0.02537 0.01895 -0.0321 0.0121 1.0000 9.000 1.2536 0.02825 0.02215 -0.0292 0.0113 1.0000 9.500 1.2785 0.03120 0.02542 -0.0261 0.0104 1.0000 10.000 1.2911 0.03457 0.02907 -0.0217 0.0098 1.0000 10.500 1.2937 0.03887 0.03367 -0.0177 0.0095 1.0000 11.000 1.2869 0.04470 0.03985 -0.0151 0.0093 1.0000 11.500 1.2694 0.05251 0.04807 -0.0147 0.0092 1.0000 12.000 1.2461 0.06207 0.05805 -0.0175 0.0092 1.0000 12.500 1.2187 0.07365 0.07002 -0.0233 0.0092 1.0000 13.000 1.1853 0.08739 0.08413 -0.0310 0.0093 1.0000 13.500 1.1401 0.10549 0.10265 -0.0419 0.0096 1.0000