XFOIL Version 6.94 Calculated polar for: AQUILA 9.3% smoothed 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 0.3909 0.00955 0.00230 -0.0653 0.6326 0.0494 0.500 0.4242 0.00706 0.00212 -0.0610 0.6139 0.8469 1.000 0.5363 0.00704 0.00199 -0.0728 0.5892 1.0000 1.500 0.5861 0.00724 0.00199 -0.0714 0.5681 1.0000 2.000 0.6359 0.00748 0.00205 -0.0701 0.5473 1.0000 2.500 0.6864 0.00770 0.00216 -0.0689 0.5260 1.0000 3.000 0.7365 0.00795 0.00230 -0.0677 0.5039 1.0000 3.500 0.7864 0.00826 0.00249 -0.0665 0.4813 1.0000 4.000 0.8367 0.00856 0.00272 -0.0654 0.4585 1.0000 4.500 0.8866 0.00891 0.00301 -0.0643 0.4345 1.0000 5.000 0.9362 0.00930 0.00333 -0.0632 0.4104 1.0000 5.500 0.9856 0.00973 0.00372 -0.0621 0.3858 1.0000 6.000 1.0343 0.01022 0.00416 -0.0609 0.3619 1.0000 6.500 1.0828 0.01072 0.00466 -0.0597 0.3379 1.0000 7.000 1.1299 0.01133 0.00522 -0.0584 0.3140 1.0000 7.500 1.1758 0.01197 0.00583 -0.0569 0.2835 1.0000 8.000 1.2220 0.01261 0.00649 -0.0555 0.2592 1.0000 8.500 1.2654 0.01339 0.00724 -0.0538 0.2301 1.0000 9.000 1.3066 0.01429 0.00811 -0.0518 0.1987 1.0000 9.500 1.3419 0.01555 0.00920 -0.0490 0.1551 1.0000 10.000 1.3699 0.01717 0.01063 -0.0453 0.1074 1.0000 10.500 1.3848 0.01919 0.01243 -0.0396 0.0659 1.0000 11.000 1.3933 0.02135 0.01452 -0.0334 0.0409 1.0000 11.500 1.3926 0.02442 0.01753 -0.0273 0.0205 1.0000 12.000 1.3939 0.02788 0.02113 -0.0229 0.0146 1.0000 12.500 1.3878 0.03256 0.02603 -0.0197 0.0124 1.0000 13.000 1.3854 0.03757 0.03130 -0.0180 0.0114 1.0000 13.500 1.3791 0.04351 0.03750 -0.0173 0.0108 1.0000 14.000 1.3686 0.05045 0.04469 -0.0176 0.0103 1.0000 14.500 1.3557 0.05816 0.05263 -0.0188 0.0099 1.0000 15.000 1.3404 0.06675 0.06145 -0.0208 0.0094 1.0000 15.500 1.3226 0.07620 0.07112 -0.0235 0.0091 1.0000 16.000 1.3074 0.08562 0.08076 -0.0264 0.0090 1.0000 16.500 1.2940 0.09505 0.09040 -0.0295 0.0088 1.0000 17.000 1.2781 0.10485 0.10045 -0.0326 0.0085 1.0000 17.500 1.2668 0.11472 0.11059 -0.0366 0.0083 1.0000 18.000 1.2561 0.12511 0.12124 -0.0417 0.0082 1.0000 18.500 1.2432 0.13621 0.13261 -0.0474 0.0081 1.0000 19.000 1.2282 0.14824 0.14491 -0.0541 0.0080 1.0000 19.500 1.2096 0.16154 0.15849 -0.0619 0.0080 1.0000 20.000 1.1877 0.17645 0.17367 -0.0711 0.0080 1.0000 20.500 1.1593 0.19427 0.19175 -0.0821 0.0082 1.0000 21.000 1.1243 0.21618 0.21387 -0.0952 0.0084 1.0000 21.500 0.8348 0.25051 0.24873 -0.1085 0.0183 1.0000 22.000 0.8415 0.26063 0.25886 -0.1117 0.0173 1.0000 22.500 0.8469 0.27057 0.26882 -0.1159 0.0172 1.0000 23.000 0.8487 0.28193 0.28018 -0.1213 0.0152 1.0000 23.500 0.8559 0.29080 0.28909 -0.1249 0.0140 1.0000 24.000 0.8648 0.29823 0.29654 -0.1279 0.0128 1.0000 24.500 0.8745 0.30737 0.30570 -0.1302 0.0122 1.0000 25.000 0.8802 0.31713 0.31549 -0.1341 0.0121 1.0000 25.500 0.8820 0.32863 0.32700 -0.1391 0.0114 1.0000 26.000 0.8884 0.33833 0.33672 -0.1428 0.0099 1.0000 26.500 0.8949 0.34671 0.34514 -0.1460 0.0090 1.0000 27.000 0.9017 0.35521 0.35367 -0.1491 0.0087 1.0000 27.500 0.9075 0.36589 0.36436 -0.1521 0.0083 1.0000 28.000 0.9110 0.37786 0.37634 -0.1564 0.0078 1.0000 28.500 0.9162 0.38803 0.38654 -0.1598 0.0071 1.0000 29.000 0.9212 0.39775 0.39629 -0.1631 0.0064 1.0000 29.500 0.9257 0.40709 0.40566 -0.1664 0.0061 1.0000 30.000 0.9316 0.41544 0.41405 -0.1687 0.0056 1.0000 30.500 0.9341 0.42677 0.42539 -0.1725 0.0056 1.0000 31.000 0.9372 0.43848 0.43713 -0.1761 0.0054 1.0000