XFOIL Version 6.96 Calculated polar for: NASA/LANGLEY LS(1)-0421 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.100 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -14.750 -0.3481 0.15969 0.15417 -0.0472 1.0000 0.1499 -14.500 -0.3627 0.15946 0.15400 -0.0442 1.0000 0.1531 -14.250 -0.6067 0.10368 0.09807 -0.0722 1.0000 0.0909 -14.000 -0.6124 0.10231 0.09676 -0.0699 1.0000 0.0915 -13.750 -0.6145 0.10160 0.09610 -0.0675 1.0000 0.0925 -13.500 -0.6053 0.10280 0.09739 -0.0642 1.0000 0.0941 -13.250 -0.9528 0.07328 0.06646 -0.0641 1.0000 0.0825 -13.000 -0.9758 0.07033 0.06332 -0.0618 0.9978 0.0824 -12.750 -0.9962 0.06735 0.06009 -0.0600 0.9937 0.0824 -12.500 -1.0109 0.06416 0.05657 -0.0586 0.9899 0.0824 -12.250 -1.0100 0.06127 0.05351 -0.0576 0.9864 0.0830 -12.000 -0.9797 0.05977 0.05216 -0.0590 0.9836 0.0847 -11.750 -0.9693 0.05759 0.04980 -0.0593 0.9800 0.0861 -11.500 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0.9420 7.750 0.8642 0.03066 0.02384 -0.0153 0.5192 0.9452 8.000 0.9049 0.02929 0.02232 -0.0161 0.4925 0.9485 8.250 0.9326 0.02898 0.02182 -0.0159 0.4599 0.9524 8.500 0.9559 0.02888 0.02143 -0.0151 0.4270 0.9568 8.750 0.9700 0.02954 0.02187 -0.0139 0.3953 0.9615 9.000 0.9862 0.03032 0.02239 -0.0131 0.3664 0.9657 9.250 0.9991 0.03127 0.02318 -0.0121 0.3409 0.9702 9.500 1.0145 0.03223 0.02396 -0.0115 0.3187 0.9753 9.750 1.0320 0.03319 0.02478 -0.0113 0.2992 0.9807 10.000 1.0511 0.03413 0.02559 -0.0113 0.2823 0.9873 10.250 1.0686 0.03481 0.02616 -0.0110 0.2683 1.0000 10.500 1.0964 0.03576 0.02689 -0.0121 0.2547 1.0000 10.750 1.1132 0.03711 0.02833 -0.0124 0.2433 1.0000 11.000 1.1394 0.03823 0.02936 -0.0135 0.2329 1.0000 11.250 1.1630 0.03938 0.03046 -0.0143 0.2233 1.0000 11.500 1.1882 0.04065 0.03174 -0.0154 0.2148 1.0000 11.750 1.2121 0.04184 0.03290 -0.0163 0.2069 1.0000 12.000 1.2397 0.04314 0.03418 -0.0176 0.1998 1.0000 12.250 1.2585 0.04459 0.03573 -0.0181 0.1932 1.0000 12.500 1.2979 0.04559 0.03652 -0.0204 0.1865 1.0000 12.750 1.3055 0.04753 0.03874 -0.0199 0.1817 1.0000 13.000 1.3278 0.04894 0.04017 -0.0208 0.1763 1.0000 13.250 1.3633 0.05026 0.04136 -0.0227 0.1710 1.0000 13.500 1.3670 0.05250 0.04390 -0.0220 0.1673 1.0000 13.750 1.3803 0.05437 0.04588 -0.0222 0.1632 1.0000 14.000 1.4297 0.05533 0.04656 -0.0252 0.1580 1.0000 14.250 1.4200 0.05816 0.04978 -0.0235 0.1556 1.0000 14.500 1.4189 0.06092 0.05281 -0.0227 0.1527 1.0000 15.000 1.4581 0.06448 0.05639 -0.0241 0.1461 1.0000 15.250 1.4590 0.06763 0.05974 -0.0236 0.1438 1.0000 15.500 1.4354 0.07189 0.06439 -0.0220 0.1423 1.0000 15.750 1.4123 0.07654 0.06937 -0.0210 0.1408 1.0000 16.000 1.3886 0.08155 0.07468 -0.0205 0.1393 1.0000 16.250 1.3652 0.08687 0.08028 -0.0206 0.1378 1.0000 16.500 1.3627 0.09052 0.08404 -0.0212 0.1359 1.0000 16.750 1.2966 0.10122 0.09519 -0.0231 0.1358 1.0000 17.000 1.2763 0.10760 0.10174 -0.0252 0.1345 1.0000 17.250 1.2536 0.11482 0.10910 -0.0280 0.1333 1.0000