XFOIL Version 6.94 Calculated polar for: NASA SC(2)-0518 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 0.3589 0.01445 0.00935 -0.0902 0.7838 0.6636 0.500 0.4254 0.01403 0.00880 -0.0925 0.7588 0.6706 1.000 0.4796 0.01406 0.00884 -0.0912 0.7330 0.6732 1.500 0.5407 0.01388 0.00856 -0.0920 0.7006 0.6776 2.000 0.6018 0.01372 0.00830 -0.0929 0.6565 0.6819 2.500 0.6543 0.01394 0.00828 -0.0917 0.5840 0.6848 3.000 0.6981 0.01515 0.00845 -0.0901 0.3940 0.6906 3.500 0.7346 0.01665 0.00928 -0.0867 0.2680 0.6934 4.000 0.7814 0.01773 0.00994 -0.0855 0.1970 0.6975 4.500 0.8309 0.01863 0.01057 -0.0849 0.1626 0.7017 5.000 0.8774 0.01947 0.01129 -0.0832 0.1425 0.7048 5.500 0.9281 0.02032 0.01199 -0.0828 0.1287 0.7095 6.000 0.9722 0.02119 0.01283 -0.0808 0.1181 0.7126 6.500 1.0148 0.02220 0.01384 -0.0787 0.1099 0.7168 7.500 1.0934 0.02424 0.01593 -0.0733 0.0979 0.7251 8.000 1.1355 0.02528 0.01699 -0.0714 0.0924 0.7306 8.500 1.1718 0.02673 0.01850 -0.0685 0.0879 0.7344 9.000 1.2134 0.02800 0.01983 -0.0668 0.0832 0.7403 10.000 1.2838 0.03116 0.02322 -0.0618 0.0747 0.7506 10.500 1.3137 0.03292 0.02502 -0.0588 0.0711 0.7552 11.500 1.3102 0.03580 0.02853 -0.0456 0.0660 0.7640 12.000 1.3335 0.03878 0.03164 -0.0432 0.0635 0.7706 12.500 1.3523 0.04205 0.03504 -0.0403 0.0614 0.7757 13.000 1.3743 0.04632 0.03944 -0.0380 0.0594 0.7818 13.500 1.3842 0.05059 0.04401 -0.0354 0.0585 0.7879 14.000 1.3900 0.05551 0.04924 -0.0331 0.0574 0.7936 14.500 1.3910 0.06130 0.05532 -0.0316 0.0563 0.8003 15.000 1.3841 0.06782 0.06216 -0.0302 0.0555 0.8051 15.500 1.3718 0.07556 0.07022 -0.0302 0.0546 0.8111 16.000 1.3493 0.08458 0.07958 -0.0310 0.0538 0.8158 16.500 1.3143 0.09580 0.09118 -0.0337 0.0532 0.8207 17.000 1.2339 0.11325 0.10915 -0.0399 0.0530 0.8216