XFOIL Version 6.94 Calculated polar for: FX 61-163 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.080 Re = 0.300 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- 0.000 0.4352 0.01432 0.00786 -0.0914 0.6254 0.7013 0.500 0.4957 0.01439 0.00783 -0.0930 0.6179 0.7091 1.000 0.5491 0.01440 0.00790 -0.0927 0.6097 0.7135 1.500 0.6058 0.01449 0.00792 -0.0931 0.6011 0.7192 2.000 0.6646 0.01460 0.00806 -0.0943 0.5930 0.7264 2.500 0.7212 0.01456 0.00806 -0.0949 0.5844 0.7313 3.000 0.7766 0.01475 0.00824 -0.0950 0.5758 0.7364 3.500 0.8312 0.01472 0.00835 -0.0951 0.5654 0.7418 4.000 0.8921 0.01483 0.00839 -0.0967 0.5551 0.7474 4.500 0.9447 0.01473 0.00849 -0.0964 0.5438 0.7523 5.000 0.9991 0.01474 0.00854 -0.0963 0.5317 0.7567 5.500 1.0520 0.01465 0.00861 -0.0961 0.5170 0.7617 6.000 1.1062 0.01470 0.00877 -0.0962 0.5011 0.7674 6.500 1.1576 0.01464 0.00881 -0.0957 0.4809 0.7718 7.000 1.2032 0.01467 0.00891 -0.0939 0.4534 0.7764 7.500 1.2471 0.01498 0.00924 -0.0921 0.4205 0.7818 8.000 1.2867 0.01562 0.00980 -0.0899 0.3775 0.7866 8.500 1.3171 0.01665 0.01066 -0.0862 0.3255 0.7917 9.000 1.3343 0.01787 0.01177 -0.0803 0.2754 0.7966 9.500 1.3414 0.01960 0.01334 -0.0733 0.2246 0.8021 10.000 1.3374 0.02220 0.01571 -0.0661 0.1725 0.8080 10.500 1.3316 0.02533 0.01874 -0.0602 0.1353 0.8142 11.000 1.3280 0.02903 0.02244 -0.0560 0.1063 0.8205 11.500 1.3247 0.03349 0.02687 -0.0533 0.0794 0.8271 12.000 1.3189 0.03868 0.03202 -0.0515 0.0561 0.8334 12.500 1.3143 0.04405 0.03743 -0.0502 0.0365 0.8407 13.000 1.3096 0.04997 0.04338 -0.0497 0.0230 0.8492 13.500 1.3016 0.05658 0.05011 -0.0496 0.0126 0.8577 14.000 1.2960 0.06358 0.05729 -0.0503 0.0074 0.8678 14.500 1.2919 0.07062 0.06458 -0.0513 0.0060 0.8806 15.000 1.2861 0.07779 0.07207 -0.0521 0.0054 0.9033