XFOIL Version 6.94 Calculated polar for: NASA SC(2)-0614 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.080 Re = 0.450 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- 0.000 0.4631 0.00956 0.00516 -0.1112 0.8291 0.6858 0.500 0.5225 0.00953 0.00499 -0.1114 0.7895 0.6909 1.000 0.5769 0.00972 0.00503 -0.1103 0.7265 0.6940 1.500 0.6199 0.01107 0.00529 -0.1075 0.4909 0.6981 2.000 0.6643 0.01270 0.00589 -0.1058 0.2816 0.7018 2.500 0.7134 0.01381 0.00645 -0.1046 0.1727 0.7047 3.000 0.7667 0.01467 0.00695 -0.1043 0.1178 0.7093 3.500 0.8193 0.01536 0.00750 -0.1036 0.0927 0.7122 4.000 0.8714 0.01607 0.00814 -0.1028 0.0808 0.7155 4.500 0.9258 0.01672 0.00875 -0.1025 0.0733 0.7203 5.500 1.0268 0.01820 0.01033 -0.1001 0.0626 0.7268 7.000 1.1705 0.02118 0.01338 -0.0956 0.0487 0.7395 7.500 1.2177 0.02192 0.01424 -0.0939 0.0445 0.7432 8.000 1.2562 0.02378 0.01612 -0.0909 0.0394 0.7478 8.500 1.2994 0.02493 0.01749 -0.0886 0.0376 0.7520 9.000 1.3371 0.02665 0.01940 -0.0854 0.0360 0.7565 9.500 1.3726 0.02852 0.02144 -0.0821 0.0349 0.7613 10.000 1.4004 0.03018 0.02329 -0.0774 0.0338 0.7655 10.500 1.4272 0.03224 0.02552 -0.0731 0.0330 0.7709 11.000 1.4501 0.03486 0.02842 -0.0685 0.0326 0.7751 11.500 1.4693 0.03790 0.03171 -0.0641 0.0322 0.7805 12.000 1.4805 0.04142 0.03555 -0.0590 0.0318 0.7846 12.500 1.4846 0.04576 0.04021 -0.0543 0.0316 0.7901 13.000 1.4787 0.05094 0.04577 -0.0496 0.0314 0.7937 13.500 1.4633 0.05743 0.05265 -0.0461 0.0312 0.7984 14.000 1.4367 0.06572 0.06135 -0.0444 0.0311 0.8020 14.500 1.3983 0.07658 0.07265 -0.0458 0.0311 0.8054 15.000 1.3394 0.09338 0.08996 -0.0540 0.0313 0.8084