XFOIL Version 6.96 Calculated polar for: GOE 372 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 ------ -------- --------- --------- -------- -------- -------- -8.750 -0.3537 0.10033 0.09810 -0.0293 1.0000 0.0183 -8.500 -0.3518 0.09746 0.09526 -0.0300 1.0000 0.0184 -8.250 -0.3510 0.09460 0.09243 -0.0303 1.0000 0.0184 -8.000 -0.2919 0.07606 0.07410 -0.0325 1.0000 0.0191 -7.750 -0.2902 0.07383 0.07190 -0.0305 1.0000 0.0197 -7.500 -0.2912 0.07264 0.07075 -0.0279 1.0000 0.0205 -7.250 -0.2823 0.06909 0.06721 -0.0302 0.9979 0.0213 -7.000 -0.2682 0.06449 0.06260 -0.0349 0.9939 0.0222 -6.750 -0.2539 0.05975 0.05787 -0.0403 0.9875 0.0232 -5.750 -0.2074 0.04545 0.04314 -0.0791 0.9741 0.0179 -5.500 -0.1639 0.01886 0.01500 -0.0991 0.9668 0.0183 -5.250 -0.1341 0.01588 0.01154 -0.1009 0.9609 0.0212 -5.000 -0.0984 0.01547 0.01103 -0.1026 0.9572 0.0241 -4.750 -0.0675 0.01467 0.00998 -0.1032 0.9498 0.0267 -4.500 -0.0367 0.01273 0.00768 -0.1043 0.9434 0.0311 -4.250 -0.0068 0.01265 0.00754 -0.1046 0.9346 0.0350 -4.000 0.0245 0.01209 0.00678 -0.1052 0.9273 0.0375 -3.750 0.0521 0.01108 0.00552 -0.1052 0.9171 0.0398 -3.500 0.0795 0.01039 0.00472 -0.1050 0.9068 0.0426 -3.250 0.1075 0.01006 0.00429 -0.1049 0.8970 0.0453 -3.000 0.1352 0.00973 0.00382 -0.1047 0.8874 0.0481 -2.750 0.1619 0.00941 0.00340 -0.1043 0.8767 0.0497 -2.500 0.1883 0.00899 0.00287 -0.1039 0.8668 0.0514 -2.250 0.2148 0.00863 0.00247 -0.1035 0.8577 0.0550 -2.000 0.2416 0.00848 0.00226 -0.1031 0.8482 0.0595 -1.750 0.2681 0.00828 0.00204 -0.1027 0.8388 0.0692 -1.500 0.2950 0.00818 0.00191 -0.1024 0.8302 0.0840 -1.250 0.3216 0.00809 0.00179 -0.1020 0.8212 0.0941 -1.000 0.3484 0.00804 0.00172 -0.1017 0.8125 0.1022 -0.750 0.3751 0.00796 0.00161 -0.1013 0.8041 0.1103 -0.500 0.4017 0.00791 0.00155 -0.1010 0.7947 0.1186 -0.250 0.4283 0.00784 0.00149 -0.1006 0.7856 0.1309 0.000 0.4546 0.00777 0.00145 -0.1002 0.7759 0.1496 0.250 0.4805 0.00763 0.00143 -0.0998 0.7646 0.1920 0.500 0.5055 0.00730 0.00147 -0.0993 0.7530 0.3285 0.750 0.5508 0.00581 0.00153 -0.1035 0.7423 1.0000 1.000 0.5767 0.00589 0.00153 -0.1030 0.7308 1.0000 1.250 0.6024 0.00597 0.00154 -0.1024 0.7175 1.0000 1.500 0.6281 0.00606 0.00156 -0.1018 0.7033 1.0000 1.750 0.6536 0.00615 0.00159 -0.1012 0.6875 1.0000 2.000 0.6786 0.00627 0.00162 -0.1005 0.6673 1.0000 2.250 0.7014 0.00646 0.00163 -0.0993 0.6251 1.0000 2.500 0.7236 0.00675 0.00170 -0.0980 0.5777 1.0000 2.750 0.7463 0.00707 0.00182 -0.0969 0.5414 1.0000 3.000 0.7690 0.00743 0.00198 -0.0959 0.5018 1.0000 3.250 0.7898 0.00794 0.00216 -0.0945 0.4309 1.0000 3.500 0.8006 0.00956 0.00275 -0.0918 0.2238 1.0000 3.750 0.8121 0.01132 0.00362 -0.0893 0.0313 1.0000 4.000 0.8364 0.01168 0.00397 -0.0885 0.0218 1.0000 4.250 0.8605 0.01204 0.00441 -0.0877 0.0196 1.0000 4.500 0.8839 0.01249 0.00495 -0.0868 0.0177 1.0000 4.750 0.9062 0.01307 0.00561 -0.0856 0.0163 1.0000 5.000 0.9269 0.01384 0.00650 -0.0841 0.0154 1.0000 5.250 0.9461 0.01474 0.00749 -0.0824 0.0151 1.0000 5.500 0.9654 0.01559 0.00841 -0.0808 0.0152 1.0000 5.750 0.9849 0.01646 0.00933 -0.0791 0.0156 1.0000 6.000 0.9999 0.01801 0.01092 -0.0769 0.0147 1.0000 6.250 1.0195 0.01927 0.01221 -0.0753 0.0147 1.0000 6.500 1.0421 0.02025 0.01325 -0.0740 0.0154 1.0000 7.000 1.0963 0.02477 0.01810 -0.0721 0.0206 1.0000 7.250 1.1192 0.02609 0.01956 -0.0710 0.0189 1.0000 7.500 1.1409 0.02779 0.02139 -0.0699 0.0174 1.0000 7.750 1.1620 0.02981 0.02350 -0.0689 0.0163 1.0000 8.000 1.1823 0.03227 0.02606 -0.0680 0.0154 1.0000 11.750 1.1339 0.07694 0.07408 -0.0333 0.0101 1.0000 12.000 1.1136 0.08257 0.07986 -0.0353 0.0102 1.0000 12.250 1.0922 0.08911 0.08655 -0.0385 0.0103 1.0000 12.500 1.0694 0.09692 0.09450 -0.0432 0.0105 1.0000 12.750 1.0444 0.10654 0.10426 -0.0497 0.0108 1.0000