XFOIL Version 6.94 Calculated polar for: GOE 430 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 ------- -------- --------- --------- -------- ------- ------- 1.000 0.7644 0.00911 0.00295 -0.1566 0.6103 0.4741 1.500 0.8189 0.00934 0.00315 -0.1561 0.5816 0.5203 2.000 0.8683 0.00979 0.00338 -0.1547 0.5141 0.5653 3.000 0.9665 0.01103 0.00417 -0.1520 0.4067 0.6476 3.500 1.0161 0.01160 0.00459 -0.1508 0.3587 0.6887 4.000 1.0474 0.01372 0.00582 -0.1470 0.1904 0.7253 4.500 1.0925 0.01429 0.00646 -0.1450 0.1544 0.7982 5.500 1.1577 0.01727 0.00883 -0.1371 0.0046 1.0000 6.000 1.2018 0.01808 0.00964 -0.1351 0.0043 1.0000 6.500 1.2407 0.01896 0.01057 -0.1321 0.0042 1.0000 7.000 1.2774 0.01999 0.01166 -0.1290 0.0042 1.0000 7.500 1.3117 0.02119 0.01294 -0.1256 0.0043 1.0000 8.000 1.3435 0.02258 0.01444 -0.1221 0.0044 1.0000 8.500 1.3725 0.02424 0.01622 -0.1185 0.0045 1.0000 9.000 1.3979 0.02623 0.01835 -0.1147 0.0047 1.0000 9.500 1.4190 0.02865 0.02093 -0.1108 0.0049 1.0000 10.000 1.4340 0.03169 0.02412 -0.1067 0.0051 1.0000 10.500 1.4422 0.03550 0.02809 -0.1026 0.0052 1.0000 11.000 1.4404 0.04048 0.03327 -0.0987 0.0054 1.0000 11.500 1.4557 0.04411 0.03703 -0.0963 0.0056 1.0000 12.000 1.4637 0.04872 0.04182 -0.0940 0.0059 1.0000 12.500 1.4614 0.05480 0.04811 -0.0921 0.0063 1.0000 13.000 1.4543 0.06179 0.05529 -0.0907 0.0066 1.0000 13.500 1.4441 0.06932 0.06298 -0.0896 0.0069 1.0000 14.000 1.4384 0.07610 0.06983 -0.0882 0.0072 1.0000 14.500 1.4491 0.08176 0.07578 -0.0883 0.0080 1.0000 15.000 1.4542 0.08665 0.08073 -0.0856 0.0089 1.0000 16.500 1.1543 0.11406 0.10952 -0.0696 0.0079 1.0000 17.000 1.1856 0.11162 0.10701 -0.0641 0.0093 1.0000