XFOIL Version 6.94 Calculated polar for: WBL FX 71-089A AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.030 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.0015 0.03105 0.01791 -0.0106 0.9989 0.4272 -2.750 0.0254 0.03095 0.01800 -0.0102 0.9989 0.4526 -2.500 0.0511 0.03079 0.01814 -0.0096 0.9989 0.4903 -2.250 0.0759 0.03057 0.01830 -0.0088 0.9989 0.5306 -2.000 0.1004 0.03034 0.01846 -0.0080 0.9989 0.5688 -1.750 0.1209 0.02983 0.01871 -0.0060 0.9989 0.6412 -1.500 0.1381 0.02907 0.01887 -0.0028 0.9989 0.7533 -1.250 0.1611 0.02834 0.01858 -0.0021 0.9989 1.0011 -1.000 0.1892 0.02892 0.01901 -0.0029 0.9989 1.0011 -0.750 0.2158 0.02956 0.01964 -0.0033 0.9989 1.0011 -0.500 0.2414 0.03032 0.02053 -0.0036 0.9989 1.0011 -0.250 0.2656 0.03154 0.02218 -0.0043 0.9989 1.0011 0.000 0.3826 0.03190 0.01998 -0.0127 0.5138 1.0011 0.250 0.4115 0.03355 0.02101 -0.0123 0.4659 1.0011 0.500 0.4410 0.03511 0.02213 -0.0122 0.4303 1.0011 0.750 0.4720 0.03669 0.02348 -0.0126 0.4043 1.0011 1.000 0.5029 0.03832 0.02493 -0.0130 0.3826 1.0011 1.250 0.5339 0.04003 0.02657 -0.0135 0.3636 1.0011 1.500 0.5660 0.04204 0.02871 -0.0144 0.3520 1.0011 1.750 0.5969 0.04424 0.03092 -0.0151 0.3430 1.0011 2.000 0.6292 0.04679 0.03377 -0.0168 0.3352 1.0011 2.250 0.6584 0.04920 0.03622 -0.0175 0.3246 1.0011 2.500 0.6883 0.05219 0.03956 -0.0191 0.3167 1.0011 2.750 0.7183 0.05579 0.04355 -0.0214 0.3146 1.0011 3.000 0.7469 0.05998 0.04813 -0.0241 0.3148 1.0011 3.250 0.7726 0.06458 0.05312 -0.0270 0.3146 1.0011 3.500 0.7943 0.06934 0.05820 -0.0297 0.3119 1.0011 3.750 0.8117 0.07536 0.06458 -0.0336 0.3156 1.0011 4.000 0.8292 0.08148 0.07097 -0.0369 0.3239 1.0011 4.250 0.8354 0.08994 0.07974 -0.0430 0.3415 1.0011 4.750 0.5424 0.12107 0.11120 -0.0779 0.8026 1.0011 5.000 0.5730 0.12650 0.11659 -0.0811 0.7746 1.0011