XFOIL Version 6.96 Calculated polar for: FX 61-163 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 = 5.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -11.250 -0.4180 0.09408 0.08919 -0.0530 1.0001 0.0230 -11.000 -0.4330 0.08602 0.08113 -0.0579 1.0001 0.0224 -10.750 -0.4541 0.07865 0.07373 -0.0624 1.0001 0.0220 -10.500 -0.4770 0.07257 0.06757 -0.0655 1.0001 0.0216 -10.250 -0.5000 0.06762 0.06254 -0.0670 1.0001 0.0213 -10.000 -0.5219 0.06365 0.05849 -0.0671 1.0001 0.0211 -9.750 -0.5438 0.06046 0.05522 -0.0658 1.0001 0.0209 -9.500 -0.5673 0.05808 0.05279 -0.0630 1.0001 0.0208 -9.250 -0.5777 0.05449 0.04902 -0.0634 0.9914 0.0206 -9.000 -0.5702 0.04954 0.04364 -0.0668 0.9639 0.0203 -8.750 -0.5548 0.04505 0.03865 -0.0696 0.9411 0.0201 -8.500 -0.5302 0.04105 0.03413 -0.0725 0.9225 0.0201 -8.250 -0.5000 0.03763 0.03024 -0.0754 0.9060 0.0202 -8.000 -0.4666 0.03475 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1.1226 0.02387 0.01620 -0.0808 0.4835 0.8020 7.250 1.1463 0.02409 0.01660 -0.0806 0.4725 0.8051 7.500 1.1680 0.02421 0.01685 -0.0797 0.4624 0.8077 7.750 1.1914 0.02427 0.01700 -0.0791 0.4525 0.8108 8.000 1.2121 0.02452 0.01740 -0.0783 0.4398 0.8142 8.250 1.2336 0.02477 0.01776 -0.0776 0.4258 0.8177 8.500 1.2531 0.02505 0.01815 -0.0766 0.4108 0.8207 8.750 1.2695 0.02539 0.01856 -0.0750 0.3951 0.8237 9.000 1.2840 0.02588 0.01914 -0.0734 0.3783 0.8273 9.250 1.2970 0.02651 0.01985 -0.0716 0.3608 0.8313 9.500 1.3078 0.02720 0.02062 -0.0696 0.3452 0.8351 9.750 1.3165 0.02802 0.02154 -0.0674 0.3293 0.8387 10.000 1.3224 0.02912 0.02268 -0.0652 0.3099 0.8427 10.250 1.3251 0.03054 0.02413 -0.0631 0.2880 0.8471 10.500 1.3232 0.03231 0.02589 -0.0609 0.2653 0.8515 10.750 1.3193 0.03443 0.02800 -0.0590 0.2424 0.8564 11.000 1.3144 0.03698 0.03049 -0.0578 0.2211 0.8615 11.250 1.3102 0.03958 0.03311 -0.0566 0.2014 0.8664 11.500 1.3053 0.04236 0.03590 -0.0557 0.1843 0.8716 11.750 1.3009 0.04532 0.03886 -0.0551 0.1689 0.8772 12.000 1.2955 0.04824 0.04184 -0.0543 0.1548 0.8830 12.250 1.2900 0.05133 0.04497 -0.0538 0.1414 0.8901 12.500 1.2836 0.05442 0.04811 -0.0531 0.1296 0.8984 12.750 1.2766 0.05760 0.05135 -0.0525 0.1190 0.9089 13.250 1.2574 0.06362 0.05754 -0.0504 0.1013 0.9896 13.500 1.2539 0.06754 0.06150 -0.0514 0.0917 0.9999 13.750 1.2474 0.07176 0.06571 -0.0523 0.0834 0.9999 14.000 1.2438 0.07575 0.06977 -0.0533 0.0745 0.9999 14.250 1.2390 0.07995 0.07405 -0.0544 0.0673 0.9999 14.500 1.2320 0.08450 0.07860 -0.0557 0.0615 0.9999 14.750 1.2282 0.08866 0.08284 -0.0568 0.0560 0.9999 15.000 1.2219 0.09327 0.08750 -0.0582 0.0510 0.9999 15.250 1.2170 0.09774 0.09203 -0.0597 0.0463 0.9999 15.500 1.2128 0.10223 0.09663 -0.0613 0.0415 0.9999 15.750 1.2058 0.10721 0.10163 -0.0632 0.0377 0.9999 16.000 1.2023 0.11174 0.10629 -0.0650 0.0337 0.9999