XFOIL Version 6.94 Calculated polar for: S4158 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 0.4347 0.01115 0.00646 -0.0923 0.7932 1.0000 0.500 0.4816 0.01115 0.00634 -0.0905 0.7850 1.0000 1.000 0.5304 0.01102 0.00612 -0.0888 0.7764 1.0000 1.500 0.5797 0.01099 0.00603 -0.0874 0.7672 1.0000 2.000 0.6309 0.01073 0.00570 -0.0859 0.7587 1.0000 2.500 0.6814 0.01062 0.00560 -0.0847 0.7469 1.0000 3.000 0.7333 0.01021 0.00517 -0.0833 0.7339 1.0000 3.500 0.7852 0.00982 0.00479 -0.0819 0.7177 1.0000 4.000 0.8371 0.00957 0.00456 -0.0808 0.6989 1.0000 4.500 0.8893 0.00938 0.00437 -0.0796 0.6766 1.0000 5.000 0.9396 0.00942 0.00439 -0.0783 0.6453 1.0000 5.500 0.9867 0.00971 0.00458 -0.0764 0.5990 1.0000 6.000 1.0292 0.01034 0.00498 -0.0739 0.5379 1.0000 6.500 1.0661 0.01126 0.00564 -0.0708 0.4706 1.0000 7.000 1.0967 0.01243 0.00654 -0.0668 0.4010 1.0000 7.500 1.1196 0.01382 0.00763 -0.0617 0.3295 1.0000 8.000 1.1310 0.01543 0.00891 -0.0550 0.2585 1.0000 8.500 1.1435 0.01734 0.01053 -0.0492 0.1971 1.0000 9.000 1.1589 0.01946 0.01236 -0.0444 0.1379 1.0000 9.500 1.1724 0.02194 0.01455 -0.0398 0.0881 1.0000 10.000 1.1877 0.02447 0.01699 -0.0356 0.0662 1.0000 10.500 1.2054 0.02693 0.01954 -0.0320 0.0538 1.0000 11.500 1.2294 0.03318 0.02598 -0.0246 0.0387 1.0000 12.000 1.2495 0.03584 0.02887 -0.0221 0.0341 1.0000 12.500 1.2601 0.03945 0.03256 -0.0194 0.0295 1.0000 14.000 1.2928 0.05174 0.04554 -0.0131 0.0216 1.0000 14.500 1.2936 0.05798 0.05214 -0.0112 0.0197 1.0000 15.000 1.2939 0.06381 0.05834 -0.0107 0.0186 1.0000 15.500 1.2872 0.07106 0.06596 -0.0110 0.0175 1.0000 16.000 1.2740 0.07967 0.07493 -0.0126 0.0168 1.0000 16.500 1.2528 0.09020 0.08585 -0.0159 0.0164 1.0000 17.000 1.2264 0.10245 0.09846 -0.0211 0.0161 1.0000 17.500 1.1888 0.11810 0.11454 -0.0292 0.0161 1.0000 18.000 1.1364 0.13910 0.13598 -0.0417 0.0165 1.0000