XFOIL Version 6.94 Calculated polar for: AG11 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 0.2408 0.00609 0.00211 -0.0408 0.9121 1.0000 0.500 0.2907 0.00621 0.00187 -0.0383 0.8494 1.0000 1.000 0.3409 0.00647 0.00174 -0.0362 0.7797 1.0000 1.500 0.3925 0.00684 0.00170 -0.0346 0.7072 1.0000 2.000 0.4451 0.00728 0.00176 -0.0335 0.6307 1.0000 2.500 0.4981 0.00779 0.00188 -0.0327 0.5499 1.0000 3.000 0.5512 0.00839 0.00209 -0.0321 0.4655 1.0000 3.500 0.6043 0.00906 0.00242 -0.0316 0.3851 1.0000 4.000 0.6575 0.00975 0.00281 -0.0313 0.3138 1.0000 4.500 0.7107 0.01050 0.00333 -0.0310 0.2520 1.0000 5.000 0.7634 0.01132 0.00393 -0.0307 0.1971 1.0000 5.500 0.8155 0.01226 0.00472 -0.0303 0.1507 1.0000 6.000 0.8669 0.01336 0.00568 -0.0298 0.1150 1.0000 6.500 0.9176 0.01454 0.00687 -0.0292 0.0930 1.0000 7.000 0.9671 0.01595 0.00831 -0.0284 0.0791 1.0000 7.500 1.0155 0.01753 0.00997 -0.0274 0.0687 1.0000 8.000 1.0644 0.01892 0.01154 -0.0265 0.0608 1.0000 8.500 1.1109 0.02094 0.01387 -0.0253 0.0542 1.0000 9.000 1.1543 0.02364 0.01674 -0.0240 0.0482 1.0000 9.500 1.1993 0.02541 0.01891 -0.0227 0.0430 1.0000 10.000 1.2370 0.02882 0.02267 -0.0211 0.0382 1.0000 10.500 1.2754 0.03087 0.02515 -0.0194 0.0334 1.0000 11.000 1.3018 0.03479 0.02957 -0.0171 0.0293 1.0000 11.500 1.3323 0.03643 0.03137 -0.0153 0.0249 1.0000 12.000 1.3422 0.04040 0.03590 -0.0123 0.0213 1.0000 13.000 1.3262 0.05101 0.04716 -0.0105 0.0173 1.0000 13.500 1.3044 0.06164 0.05823 -0.0171 0.0163 1.0000 14.000 1.2578 0.07896 0.07597 -0.0286 0.0168 1.0000 14.500 1.1743 0.10571 0.10317 -0.0447 0.0198 1.0000