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 = 0.200 e 6 Ncrit = 5.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -8.250 -0.6428 0.09051 0.08711 0.0075 1.0000 0.0157 -8.000 -0.6444 0.08626 0.08290 0.0047 1.0000 0.0158 -7.750 -0.6479 0.08196 0.07864 0.0012 1.0000 0.0157 -7.500 -0.6471 0.07724 0.07391 -0.0027 1.0000 0.0158 -7.250 -0.6438 0.07238 0.06899 -0.0060 1.0000 0.0161 -7.000 -0.6379 0.06769 0.06422 -0.0086 1.0000 0.0163 -6.750 -0.6297 0.06299 0.05940 -0.0106 1.0000 0.0169 -6.250 -0.5967 0.05565 0.05145 -0.0124 1.0000 0.0190 -5.750 -0.5680 0.04805 0.04334 -0.0117 1.0000 0.0191 -5.250 -0.5418 0.03666 0.03155 -0.0102 1.0000 0.0088 -5.000 -0.5243 0.03273 0.02726 -0.0090 1.0000 0.0081 -4.750 -0.5044 0.02908 0.02321 -0.0075 1.0000 0.0074 -4.500 -0.4828 0.02555 0.01922 -0.0058 1.0000 0.0068 -4.250 -0.4598 0.02230 0.01547 -0.0039 1.0000 0.0065 -4.000 -0.4363 0.01995 0.01273 -0.0026 1.0000 0.0067 -3.750 -0.4120 0.01881 0.01135 -0.0016 1.0000 0.0090 -3.500 -0.3877 0.01693 0.00914 -0.0003 1.0000 0.0094 -3.250 -0.3640 0.01529 0.00729 0.0010 1.0000 0.0098 -3.000 -0.3426 0.01352 0.00538 0.0027 1.0000 0.0107 -2.750 -0.3208 0.01243 0.00420 0.0039 1.0000 0.0149 -2.500 -0.2977 0.01168 0.00327 0.0051 1.0000 0.0192 -2.250 -0.2750 0.01090 0.00251 0.0061 1.0000 0.0536 -2.000 -0.2622 0.00867 0.00204 0.0079 1.0000 0.4695 -1.750 -0.2474 0.00770 0.00191 0.0108 1.0000 0.6852 -1.500 -0.2257 0.00721 0.00192 0.0131 1.0000 0.8454 -1.250 -0.1447 0.00731 0.00192 0.0026 1.0000 0.9724 -1.000 -0.1018 0.00730 0.00178 -0.0009 1.0000 0.9880 -0.750 -0.0626 0.00726 0.00166 -0.0036 1.0000 0.9982 -0.500 -0.0392 0.00723 0.00156 -0.0030 1.0000 1.0000 -0.250 -0.0197 0.00720 0.00151 -0.0015 1.0000 1.0000 0.000 0.0000 0.00719 0.00149 0.0000 1.0000 1.0000 0.250 0.0197 0.00720 0.00151 0.0015 1.0000 1.0000 0.500 0.0392 0.00723 0.00156 0.0030 1.0000 1.0000 0.750 0.0627 0.00727 0.00166 0.0036 0.9981 1.0000 1.000 0.1019 0.00730 0.00179 0.0008 0.9880 1.0000 1.250 0.1460 0.00731 0.00192 -0.0029 0.9716 1.0000 1.500 0.2258 0.00721 0.00192 -0.0132 0.8462 1.0000 1.750 0.2477 0.00770 0.00192 -0.0109 0.6856 1.0000 2.000 0.2622 0.00869 0.00204 -0.0079 0.4647 1.0000 2.250 0.2752 0.01090 0.00251 -0.0062 0.0535 1.0000 2.500 0.2980 0.01167 0.00326 -0.0051 0.0193 1.0000 2.750 0.3211 0.01242 0.00419 -0.0040 0.0143 1.0000 3.000 0.3428 0.01350 0.00536 -0.0027 0.0107 1.0000 3.250 0.3641 0.01529 0.00729 -0.0011 0.0099 1.0000 3.500 0.3878 0.01694 0.00916 0.0003 0.0094 1.0000 3.750 0.4123 0.01869 0.01123 0.0016 0.0086 1.0000 4.000 0.4365 0.02001 0.01281 0.0026 0.0068 1.0000 4.250 0.4599 0.02230 0.01547 0.0039 0.0065 1.0000 4.500 0.4828 0.02557 0.01924 0.0058 0.0069 1.0000 4.750 0.5044 0.02906 0.02318 0.0075 0.0074 1.0000 5.000 0.5243 0.03273 0.02726 0.0090 0.0081 1.0000 5.250 0.5417 0.03663 0.03152 0.0102 0.0089 1.0000 5.750 0.5678 0.04803 0.04332 0.0118 0.0191 1.0000 6.250 0.5964 0.05562 0.05142 0.0125 0.0190 1.0000 6.500 0.6098 0.05939 0.05547 0.0123 0.0188 1.0000 6.750 0.6290 0.06299 0.05939 0.0107 0.0171 1.0000 7.000 0.6376 0.06760 0.06413 0.0087 0.0164 1.0000 7.250 0.6433 0.07239 0.06900 0.0061 0.0161 1.0000 7.500 0.6469 0.07714 0.07381 0.0028 0.0159 1.0000 7.750 0.6477 0.08193 0.07861 -0.0011 0.0158 1.0000 8.000 0.6441 0.08627 0.08290 -0.0047 0.0158 1.0000 8.250 0.6427 0.09049 0.08708 -0.0075 0.0155 1.0000