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 = 0.450 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- 0.500 0.2813 0.00538 0.00139 -0.0354 0.7733 1.0000 1.000 0.3344 0.00571 0.00132 -0.0343 0.7027 1.0000 1.500 0.3882 0.00609 0.00132 -0.0336 0.6307 1.0000 2.000 0.4424 0.00653 0.00139 -0.0330 0.5559 1.0000 2.500 0.4965 0.00703 0.00153 -0.0326 0.4761 1.0000 3.000 0.5506 0.00760 0.00175 -0.0323 0.3982 1.0000 3.500 0.6048 0.00819 0.00203 -0.0321 0.3281 1.0000 4.000 0.6589 0.00880 0.00240 -0.0319 0.2674 1.0000 4.500 0.7128 0.00943 0.00284 -0.0317 0.2155 1.0000 5.000 0.7663 0.01013 0.00334 -0.0315 0.1683 1.0000 5.500 0.8195 0.01090 0.00398 -0.0312 0.1288 1.0000 6.000 0.8720 0.01179 0.00475 -0.0308 0.0983 1.0000 6.500 0.9238 0.01278 0.00570 -0.0303 0.0793 1.0000 7.000 0.9753 0.01376 0.00677 -0.0298 0.0675 1.0000 7.500 1.0256 0.01492 0.00802 -0.0291 0.0586 1.0000 8.000 1.0736 0.01645 0.00964 -0.0281 0.0511 1.0000 8.500 1.1238 0.01743 0.01083 -0.0274 0.0457 1.0000 9.000 1.1696 0.01921 0.01282 -0.0262 0.0406 1.0000 9.500 1.2159 0.02066 0.01443 -0.0252 0.0357 1.0000 10.000 1.2591 0.02261 0.01672 -0.0238 0.0315 1.0000 10.500 1.2982 0.02488 0.01917 -0.0223 0.0270 1.0000 11.000 1.3452 0.02538 0.01990 -0.0217 0.0225 1.0000 11.500 1.3832 0.02726 0.02209 -0.0202 0.0183 1.0000 12.000 1.4137 0.02982 0.02493 -0.0182 0.0148 1.0000 12.500 1.4329 0.03313 0.02852 -0.0155 0.0118 1.0000 13.000 1.4396 0.03672 0.03247 -0.0120 0.0098 1.0000 13.500 1.4150 0.04449 0.04056 -0.0120 0.0090 1.0000 14.000 1.3897 0.05617 0.05271 -0.0191 0.0086 1.0000 14.500 1.3521 0.07203 0.06892 -0.0295 0.0087 1.0000 15.500 1.2642 0.10610 0.10351 -0.0478 0.0093 1.0000 16.000 1.2057 0.12697 0.12470 -0.0586 0.0102 1.0000 16.500 1.1477 0.15012 0.14805 -0.0711 0.0115 1.0000