XFOIL Version 6.94 Calculated polar for: FX S 02/1-158 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 ------- -------- --------- --------- -------- ------- ------- 0.500 0.5076 0.01051 0.00292 -0.0917 0.4488 0.4982 1.000 0.5641 0.01057 0.00292 -0.0918 0.4405 0.5076 1.500 0.6205 0.01070 0.00306 -0.0920 0.4331 0.5166 2.000 0.6778 0.01072 0.00311 -0.0924 0.4281 0.5214 2.500 0.7337 0.01077 0.00303 -0.0925 0.4097 0.5265 3.000 0.7897 0.01099 0.00321 -0.0926 0.4020 0.5316 3.500 0.8464 0.01103 0.00336 -0.0929 0.4000 0.5405 4.000 0.9016 0.01113 0.00342 -0.0930 0.3727 0.5508 4.500 0.9267 0.01441 0.00535 -0.0896 0.1220 0.5530 5.000 0.9763 0.01513 0.00601 -0.0890 0.0952 0.5619 5.500 1.0214 0.01617 0.00696 -0.0879 0.0500 0.5730 6.000 1.0671 0.01705 0.00788 -0.0868 0.0257 0.5834 6.500 1.1150 0.01770 0.00863 -0.0860 0.0245 0.5963 7.000 1.1548 0.01885 0.00990 -0.0842 0.0042 0.6050 7.500 1.1965 0.01973 0.01093 -0.0826 0.0040 0.6234 8.000 1.2308 0.02093 0.01238 -0.0801 0.0040 0.6463 8.500 1.2551 0.02233 0.01406 -0.0763 0.0041 0.6756 9.000 1.2675 0.02456 0.01681 -0.0723 0.0042 0.7612 10.500 1.2608 0.04133 0.03469 -0.0700 0.0047 1.0000 11.000 1.2546 0.04773 0.04127 -0.0699 0.0043 1.0000 11.500 1.2404 0.05511 0.04884 -0.0696 0.0046 1.0000 12.000 1.2332 0.06185 0.05576 -0.0697 0.0044 1.0000 12.500 1.2319 0.06809 0.06213 -0.0696 0.0049 1.0000 13.000 1.2314 0.07444 0.06864 -0.0701 0.0048 1.0000 13.500 1.2287 0.08006 0.07453 -0.0671 0.0072 1.0000 14.000 1.2412 0.08446 0.07908 -0.0664 0.0073 1.0000 14.500 1.2551 0.08850 0.08332 -0.0652 0.0073 1.0000