XFOIL Version 6.96 Calculated polar for: GOE 443 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.500 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -8.500 -0.5624 0.08080 0.07868 0.0041 1.0000 0.0104 -8.250 -0.5663 0.07600 0.07390 0.0021 1.0000 0.0105 -8.000 -0.5722 0.07116 0.06909 -0.0002 1.0000 0.0106 -7.750 -0.5819 0.06628 0.06423 -0.0033 1.0000 0.0105 -7.500 -0.5958 0.06127 0.05922 -0.0069 1.0000 0.0103 -7.250 -0.5999 0.05589 0.05377 -0.0096 1.0000 0.0103 -7.000 -0.6001 0.05060 0.04840 -0.0115 1.0000 0.0105 -6.750 -0.5967 0.04545 0.04313 -0.0128 1.0000 0.0107 -6.500 -0.5899 0.04048 0.03802 -0.0135 1.0000 0.0111 -6.250 -0.5805 0.03574 0.03311 -0.0137 1.0000 0.0116 -6.000 -0.5681 0.03128 0.02845 -0.0136 1.0000 0.0124 -5.750 -0.5514 0.02750 0.02444 -0.0129 1.0000 0.0138 -5.500 -0.5262 0.02632 0.02303 -0.0116 1.0000 0.0152 -3.750 -0.4206 0.01570 0.00977 0.0006 1.0000 0.0169 -3.500 -0.3958 0.01464 0.00861 0.0018 1.0000 0.0149 -3.250 -0.3731 0.01257 0.00636 0.0033 1.0000 0.0145 -3.000 -0.3508 0.01124 0.00492 0.0047 1.0000 0.0147 -2.750 -0.3296 0.01004 0.00361 0.0063 1.0000 0.0164 -2.500 -0.3079 0.00915 0.00265 0.0076 1.0000 0.0206 -2.250 -0.2838 0.00894 0.00242 0.0084 1.0000 0.0265 -2.000 -0.2614 0.00822 0.00168 0.0096 1.0000 0.0523 -1.750 -0.2433 0.00683 0.00128 0.0108 1.0000 0.3292 -1.500 -0.2302 0.00545 0.00120 0.0134 1.0000 0.6654 -1.250 -0.2159 0.00486 0.00127 0.0167 1.0000 0.8353 -1.000 -0.1639 0.00476 0.00142 0.0121 1.0000 0.9552 -0.750 -0.0898 0.00488 0.00148 0.0021 1.0000 0.9898 -0.500 -0.0340 0.00492 0.00145 -0.0042 1.0000 1.0000 -0.250 -0.0172 0.00488 0.00140 -0.0021 1.0000 1.0000 0.000 0.0000 0.00487 0.00138 0.0000 1.0000 1.0000 0.250 0.0172 0.00488 0.00140 0.0021 1.0000 1.0000 0.500 0.0341 0.00492 0.00145 0.0041 1.0000 1.0000 0.750 0.0892 0.00488 0.00148 -0.0019 0.9901 1.0000 1.000 0.1629 0.00476 0.00142 -0.0119 0.9557 1.0000 1.250 0.2156 0.00485 0.00127 -0.0167 0.8368 1.0000 1.500 0.2308 0.00538 0.00121 -0.0135 0.6831 1.0000 1.750 0.2428 0.00683 0.00127 -0.0107 0.3290 1.0000 2.000 0.2609 0.00822 0.00168 -0.0095 0.0528 1.0000 2.250 0.2833 0.00894 0.00242 -0.0083 0.0266 1.0000 2.500 0.3075 0.00915 0.00265 -0.0075 0.0209 1.0000 2.750 0.3290 0.01009 0.00366 -0.0061 0.0162 1.0000 3.000 0.3504 0.01123 0.00491 -0.0046 0.0147 1.0000 3.250 0.3727 0.01254 0.00633 -0.0032 0.0145 1.0000 3.500 0.3954 0.01458 0.00855 -0.0017 0.0148 1.0000 3.750 0.4200 0.01570 0.00977 -0.0005 0.0169 1.0000 6.000 0.5684 0.03120 0.02838 0.0135 0.0124 1.0000 6.250 0.5806 0.03569 0.03307 0.0137 0.0116 1.0000 6.500 0.5902 0.04044 0.03798 0.0134 0.0111 1.0000 6.750 0.5969 0.04542 0.04311 0.0127 0.0107 1.0000 7.000 0.6004 0.05056 0.04835 0.0114 0.0105 1.0000 7.250 0.6002 0.05585 0.05374 0.0095 0.0103 1.0000 7.500 0.5960 0.06124 0.05918 0.0068 0.0103 1.0000 7.750 0.5822 0.06624 0.06419 0.0031 0.0105 1.0000 8.000 0.5726 0.07111 0.06904 0.0000 0.0107 1.0000 8.250 0.5666 0.07597 0.07387 -0.0023 0.0105 1.0000 8.500 0.5627 0.08074 0.07863 -0.0043 0.0104 1.0000