XFOIL Version 6.96 Calculated polar for: RAE6-9CK AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 1.000 e 6 Ncrit = 5.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -19.000 -0.6753 0.20875 0.20691 0.0137 1.0000 0.0039 -18.500 -0.8211 0.15006 0.14813 -0.0093 1.0000 0.0051 -18.250 -0.9366 0.11929 0.11708 -0.0244 1.0000 0.0049 -18.000 -0.9916 0.10356 0.10114 -0.0323 1.0000 0.0048 -17.750 -1.0368 0.09036 0.08774 -0.0393 1.0000 0.0047 -17.500 -1.0690 0.08041 0.07760 -0.0448 1.0000 0.0047 -17.250 -1.0883 0.07341 0.07045 -0.0485 1.0000 0.0047 -17.000 -1.1021 0.06763 0.06455 -0.0514 1.0000 0.0047 -16.750 -1.1126 0.06268 0.05948 -0.0537 1.0000 0.0048 -16.500 -1.1237 0.05788 0.05455 -0.0558 1.0000 0.0048 -16.250 -1.1287 0.05414 0.05071 -0.0572 1.0000 0.0048 -16.000 -1.1367 0.05024 0.04668 -0.0585 1.0000 0.0048 -15.750 -1.1378 0.04735 0.04370 -0.0593 1.0000 0.0049 -15.500 -1.1383 0.04464 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5.000 0.7152 0.01113 0.00396 -0.0561 0.0237 0.6282 5.250 0.7411 0.01133 0.00419 -0.0558 0.0219 0.6330 5.500 0.7664 0.01158 0.00444 -0.0554 0.0198 0.6375 5.750 0.7914 0.01185 0.00474 -0.0549 0.0179 0.6426 6.000 0.8167 0.01208 0.00501 -0.0546 0.0169 0.6480 6.250 0.8417 0.01234 0.00530 -0.0541 0.0157 0.6534 6.500 0.8662 0.01263 0.00562 -0.0536 0.0143 0.6595 7.000 0.9150 0.01324 0.00631 -0.0525 0.0126 0.6711 7.250 0.9391 0.01355 0.00666 -0.0520 0.0120 0.6773 7.500 0.9630 0.01387 0.00704 -0.0514 0.0113 0.6833 7.750 0.9863 0.01424 0.00744 -0.0507 0.0107 0.6906 8.000 1.0087 0.01471 0.00796 -0.0499 0.0099 0.6974 8.250 1.0318 0.01507 0.00839 -0.0492 0.0097 0.7049 8.500 1.0548 0.01544 0.00883 -0.0484 0.0093 0.7118 8.750 1.0774 0.01582 0.00929 -0.0477 0.0089 0.7195 9.000 1.0998 0.01621 0.00973 -0.0469 0.0085 0.7273 9.250 1.1217 0.01663 0.01022 -0.0460 0.0081 0.7363 9.500 1.1427 0.01711 0.01077 -0.0450 0.0078 0.7453 10.000 1.1815 0.01834 0.01217 -0.0425 0.0071 0.7646 10.250 1.2015 0.01883 0.01276 -0.0414 0.0070 0.7751 10.500 1.2206 0.01937 0.01341 -0.0401 0.0069 0.7870 10.750 1.2383 0.01991 0.01406 -0.0386 0.0067 0.8011 11.000 1.2539 0.02050 0.01476 -0.0367 0.0065 0.8168 11.250 1.2686 0.02110 0.01550 -0.0347 0.0064 0.8350 11.500 1.2821 0.02172 0.01626 -0.0325 0.0062 0.8589 11.750 1.2938 0.02219 0.01692 -0.0300 0.0061 0.9020 12.000 1.3072 0.02286 0.01779 -0.0280 0.0060 1.0000 12.250 1.3213 0.02359 0.01858 -0.0262 0.0058 1.0000 12.500 1.3336 0.02444 0.01951 -0.0243 0.0057 1.0000 12.750 1.3456 0.02532 0.02045 -0.0225 0.0055 1.0000 13.000 1.3562 0.02629 0.02151 -0.0206 0.0054 1.0000 13.250 1.3649 0.02743 0.02274 -0.0185 0.0053 1.0000 13.500 1.3719 0.02870 0.02411 -0.0165 0.0052 1.0000 13.750 1.3751 0.03031 0.02584 -0.0142 0.0051 1.0000 14.000 1.3772 0.03203 0.02769 -0.0121 0.0050 1.0000 14.250 1.3780 0.03392 0.02972 -0.0101 0.0049 1.0000 14.500 1.3822 0.03558 0.03149 -0.0086 0.0049 1.0000 14.750 1.3853 0.03739 0.03343 -0.0072 0.0048 1.0000 15.000 1.3851 0.03960 0.03576 -0.0060 0.0048 1.0000 15.250 1.3836 0.04202 0.03832 -0.0049 0.0048 1.0000 15.500 1.3828 0.04451 0.04093 -0.0043 0.0047 1.0000 15.750 1.3759 0.04779 0.04436 -0.0039 0.0047 1.0000 16.000 1.3678 0.05144 0.04815 -0.0040 0.0047 1.0000 16.250 1.3627 0.05497 0.05180 -0.0046 0.0046 1.0000 16.500 1.3467 0.06024 0.05725 -0.0061 0.0046 1.0000 16.750 1.3327 0.06574 0.06291 -0.0083 0.0045 1.0000 17.000 1.3124 0.07290 0.07023 -0.0120 0.0046 1.0000 17.250 1.2860 0.08216 0.07969 -0.0175 0.0045 1.0000 17.500 1.2442 0.09615 0.09392 -0.0266 0.0046 1.0000