XFOIL Version 6.96 Calculated polar for: NASA/AMES A-01 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.200 e 6 Ncrit = 5.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -9.750 -0.7277 0.06507 0.06155 -0.0217 1.0000 0.0207 -9.500 -0.7634 0.05562 0.05177 -0.0234 1.0000 0.0207 -9.250 -0.8094 0.04179 0.03688 -0.0222 1.0000 0.0213 -9.000 -0.7946 0.04004 0.03502 -0.0219 1.0000 0.0218 -8.750 -0.7755 0.03898 0.03386 -0.0217 1.0000 0.0224 -8.500 -0.7595 0.03688 0.03152 -0.0212 1.0000 0.0230 -8.250 -0.7419 0.03493 0.02934 -0.0208 1.0000 0.0238 -8.000 -0.7262 0.03211 0.02612 -0.0199 1.0000 0.0250 -7.750 -0.7108 0.02896 0.02242 -0.0187 1.0000 0.0261 -7.500 -0.6902 0.02653 0.01949 -0.0180 0.9942 0.0270 -7.250 -0.6636 0.02489 0.01769 -0.0185 0.9876 0.0278 -7.000 -0.6345 0.02403 0.01674 -0.0193 0.9828 0.0288 -6.750 -0.6060 0.02327 0.01586 -0.0199 0.9780 0.0301 -6.500 -0.5769 0.02226 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-0.0113 0.2711 1.0000 3.500 0.5048 0.01274 0.00424 -0.0105 0.2566 1.0000 3.750 0.5298 0.01298 0.00442 -0.0100 0.2504 1.0000 4.000 0.5549 0.01322 0.00462 -0.0096 0.2446 1.0000 4.250 0.5802 0.01344 0.00484 -0.0092 0.2386 1.0000 4.500 0.6050 0.01372 0.00507 -0.0087 0.2335 1.0000 4.750 0.6302 0.01397 0.00533 -0.0083 0.2291 1.0000 5.000 0.6554 0.01421 0.00559 -0.0078 0.2240 1.0000 5.250 0.6803 0.01451 0.00586 -0.0074 0.2191 1.0000 5.500 0.7053 0.01481 0.00616 -0.0070 0.2148 1.0000 5.750 0.7306 0.01506 0.00647 -0.0066 0.2099 1.0000 6.000 0.7556 0.01535 0.00677 -0.0062 0.2050 1.0000 6.250 0.7805 0.01569 0.00710 -0.0058 0.2003 1.0000 6.500 0.8060 0.01593 0.00743 -0.0054 0.1948 1.0000 6.750 0.8311 0.01622 0.00775 -0.0050 0.1896 1.0000 7.000 0.8561 0.01653 0.00809 -0.0047 0.1841 1.0000 7.250 0.8815 0.01677 0.00842 -0.0043 0.1776 1.0000 7.500 0.9061 0.01710 0.00875 -0.0040 0.1718 1.0000 7.750 0.9315 0.01735 0.00912 -0.0037 0.1648 1.0000 8.000 0.9560 0.01765 0.00943 -0.0033 0.1578 1.0000 8.250 0.9810 0.01791 0.00979 -0.0030 0.1488 1.0000 8.500 1.0055 0.01824 0.01017 -0.0026 0.1395 1.0000 8.750 1.0295 0.01861 0.01056 -0.0023 0.1295 1.0000 9.000 1.0529 0.01905 0.01101 -0.0018 0.1185 1.0000 9.250 1.0761 0.01954 0.01153 -0.0014 0.1062 1.0000 9.500 1.0982 0.02015 0.01215 -0.0009 0.0929 1.0000 9.750 1.1194 0.02087 0.01285 -0.0002 0.0798 1.0000 10.000 1.1395 0.02171 0.01368 0.0005 0.0685 1.0000 10.500 1.1773 0.02353 0.01557 0.0022 0.0514 1.0000 10.750 1.1947 0.02451 0.01662 0.0032 0.0458 1.0000 11.000 1.2109 0.02555 0.01773 0.0043 0.0413 1.0000 11.250 1.2249 0.02671 0.01896 0.0056 0.0380 1.0000 11.500 1.2392 0.02775 0.02014 0.0069 0.0352 1.0000 11.750 1.2504 0.02896 0.02144 0.0085 0.0328 1.0000 12.000 1.2549 0.03042 0.02297 0.0107 0.0309 1.0000 12.250 1.2610 0.03173 0.02444 0.0128 0.0296 1.0000 12.500 1.2663 0.03321 0.02606 0.0145 0.0283 1.0000 12.750 1.2698 0.03491 0.02790 0.0161 0.0272 1.0000 13.000 1.2718 0.03684 0.02996 0.0174 0.0263 1.0000 13.250 1.2716 0.03908 0.03232 0.0184 0.0255 1.0000 13.500 1.2685 0.04177 0.03512 0.0191 0.0248 1.0000 13.750 1.2618 0.04505 0.03853 0.0192 0.0242 1.0000 14.000 1.2542 0.04873 0.04233 0.0187 0.0237 1.0000 14.250 1.2491 0.05249 0.04628 0.0177 0.0233 1.0000 14.500 1.2414 0.05690 0.05087 0.0159 0.0229 1.0000 14.750 1.2316 0.06203 0.05617 0.0135 0.0226 1.0000 15.000 1.2198 0.06779 0.06212 0.0104 0.0223 1.0000 15.250 1.2064 0.07406 0.06854 0.0071 0.0220 1.0000 15.500 1.1920 0.08057 0.07522 0.0036 0.0218 1.0000 15.750 1.1774 0.08714 0.08192 0.0003 0.0216 1.0000