XFOIL Version 6.94 Calculated polar for: RG 8 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 0.3806 0.00679 0.00212 -0.0935 0.8333 0.6991 0.500 0.4357 0.00686 0.00211 -0.0929 0.7976 0.7193 1.000 0.4891 0.00698 0.00214 -0.0920 0.7592 0.7398 1.500 0.5408 0.00715 0.00222 -0.0908 0.7177 0.7600 2.000 0.5912 0.00734 0.00233 -0.0894 0.6738 0.7809 2.500 0.6402 0.00757 0.00250 -0.0877 0.6277 0.8045 3.000 0.6870 0.00783 0.00270 -0.0856 0.5801 0.8340 3.500 0.7304 0.00812 0.00295 -0.0827 0.5246 0.8742 4.000 0.7656 0.00857 0.00316 -0.0780 0.4365 0.9552 5.000 0.8640 0.01022 0.00401 -0.0764 0.2736 1.0000 5.500 0.9107 0.01107 0.00458 -0.0750 0.2128 1.0000 6.000 0.9518 0.01238 0.00537 -0.0728 0.1250 1.0000 6.500 0.9939 0.01357 0.00625 -0.0707 0.0702 1.0000 7.000 1.0356 0.01474 0.00721 -0.0684 0.0341 1.0000 7.500 1.0760 0.01594 0.00832 -0.0659 0.0128 1.0000 8.000 1.1149 0.01718 0.00958 -0.0631 0.0043 1.0000 8.500 1.1540 0.01829 0.01083 -0.0603 0.0039 1.0000 9.000 1.1882 0.01951 0.01223 -0.0567 0.0036 1.0000 9.500 1.2165 0.02091 0.01383 -0.0521 0.0035 1.0000 10.000 1.2403 0.02256 0.01568 -0.0471 0.0034 1.0000 10.500 1.2588 0.02453 0.01787 -0.0418 0.0034 1.0000 11.000 1.2714 0.02688 0.02045 -0.0362 0.0034 1.0000 11.500 1.2773 0.02980 0.02362 -0.0307 0.0034 1.0000 12.000 1.2782 0.03337 0.02745 -0.0258 0.0034 1.0000 12.500 1.2735 0.03788 0.03223 -0.0219 0.0034 1.0000 13.000 1.2664 0.04328 0.03794 -0.0195 0.0035 1.0000 13.500 1.2547 0.05001 0.04498 -0.0187 0.0035 1.0000 14.000 1.2399 0.05803 0.05334 -0.0198 0.0036 1.0000 14.500 1.2204 0.06781 0.06347 -0.0229 0.0037 1.0000 15.000 1.1973 0.07940 0.07542 -0.0280 0.0037 1.0000 15.500 1.1708 0.09290 0.08928 -0.0351 0.0038 1.0000 16.000 1.1433 0.10792 0.10462 -0.0438 0.0039 1.0000 16.500 1.1131 0.12474 0.12176 -0.0540 0.0039 1.0000 17.000 1.0831 0.14264 0.13994 -0.0650 0.0040 1.0000 17.500 1.0525 0.16185 0.15940 -0.0768 0.0041 1.0000 18.000 1.0115 0.18614 0.18389 -0.0907 0.0043 1.0000 20.000 0.8374 0.26597 0.26430 -0.1329 0.0116 1.0000 20.500 0.8468 0.27415 0.27249 -0.1365 0.0110 1.0000 21.000 0.8585 0.28135 0.27972 -0.1391 0.0107 1.0000 21.500 0.8655 0.29119 0.28957 -0.1432 0.0106 1.0000 22.000 0.8708 0.30271 0.30109 -0.1481 0.0100 1.0000 22.500 0.8790 0.31201 0.31041 -0.1518 0.0094 1.0000