XFOIL Version 6.96 Calculated polar for: GOE 392 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 = 5.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -8.750 -0.2187 0.09436 0.09217 -0.0602 0.9835 0.0059 -8.500 -0.2056 0.09018 0.08801 -0.0646 0.9795 0.0062 -8.250 -0.1955 0.08644 0.08429 -0.0681 0.9721 0.0065 -8.000 -0.1855 0.08283 0.08068 -0.0709 0.9647 0.0065 -7.500 -0.1599 0.07464 0.07249 -0.0792 0.9446 0.0065 -5.500 0.0158 0.03206 0.02855 -0.1323 0.8040 0.0059 -5.250 0.0329 0.02849 0.02465 -0.1327 0.7939 0.0054 -5.000 0.0519 0.02492 0.02071 -0.1323 0.7846 0.0050 -4.750 0.0720 0.02116 0.01646 -0.1313 0.7763 0.0046 -4.500 0.0931 0.01720 0.01192 -0.1298 0.7681 0.0042 -4.250 0.1163 0.01516 0.00949 -0.1289 0.7604 0.0044 -4.000 0.1404 0.01369 0.00771 -0.1280 0.7522 0.0046 -3.750 0.1650 0.01258 0.00634 -0.1273 0.7443 0.0051 -3.500 0.1899 0.01182 0.00538 -0.1267 0.7363 0.0060 -3.250 0.2145 0.01107 0.00448 -0.1261 0.7283 0.0071 -3.000 0.2389 0.01043 0.00369 -0.1254 0.7201 0.0080 -2.750 0.2643 0.01002 0.00317 -0.1250 0.7114 0.0092 -2.500 0.2901 0.00981 0.00285 -0.1246 0.7029 0.0115 -2.250 0.3151 0.00942 0.00234 -0.1240 0.6929 0.0172 -2.000 0.3407 0.00930 0.00212 -0.1236 0.6804 0.0244 -1.750 0.3661 0.00922 0.00191 -0.1232 0.6669 0.0313 -1.500 0.3915 0.00910 0.00170 -0.1228 0.6544 0.0403 -1.250 0.4161 0.00876 0.00153 -0.1224 0.6440 0.1297 -1.000 0.4395 0.00830 0.00150 -0.1219 0.6341 0.3001 -0.750 0.4650 0.00826 0.00149 -0.1216 0.6234 0.3521 -0.500 0.4909 0.00826 0.00148 -0.1213 0.6131 0.3770 -0.250 0.5165 0.00827 0.00147 -0.1210 0.6037 0.4013 0.000 0.5420 0.00829 0.00146 -0.1206 0.5937 0.4209 0.250 0.5676 0.00828 0.00147 -0.1203 0.5835 0.4418 0.750 0.6140 0.00788 0.00161 -0.1188 0.5658 0.6560 1.000 0.6346 0.00762 0.00170 -0.1173 0.5580 0.7792 1.250 0.6888 0.00742 0.00178 -0.1234 0.5487 1.0000 1.500 0.7142 0.00753 0.00185 -0.1230 0.5404 1.0000 1.750 0.7394 0.00765 0.00193 -0.1226 0.5326 1.0000 2.000 0.7646 0.00777 0.00201 -0.1222 0.5245 1.0000 2.250 0.7899 0.00789 0.00211 -0.1219 0.5172 1.0000 2.500 0.8152 0.00802 0.00224 -0.1215 0.5099 1.0000 2.750 0.8395 0.00818 0.00237 -0.1209 0.4966 1.0000 3.000 0.8595 0.00854 0.00253 -0.1195 0.4557 1.0000 3.250 0.8787 0.00897 0.00273 -0.1180 0.4088 1.0000 3.500 0.8908 0.00993 0.00319 -0.1153 0.3167 1.0000 3.750 0.8802 0.01268 0.00462 -0.1088 0.0227 1.0000 4.000 0.9024 0.01301 0.00499 -0.1079 0.0140 1.0000 4.250 0.9244 0.01337 0.00543 -0.1068 0.0104 1.0000 4.500 0.9441 0.01389 0.00607 -0.1054 0.0080 1.0000 4.750 0.9635 0.01440 0.00669 -0.1038 0.0069 1.0000 5.000 0.9822 0.01493 0.00727 -0.1023 0.0060 1.0000 5.250 0.9980 0.01561 0.00800 -0.1002 0.0049 1.0000 5.500 1.0118 0.01638 0.00886 -0.0977 0.0047 1.0000 5.750 1.0218 0.01719 0.00976 -0.0944 0.0042 1.0000 6.000 1.0285 0.01813 0.01079 -0.0906 0.0041 1.0000 6.250 1.0336 0.01920 0.01197 -0.0867 0.0039 1.0000 6.500 1.0391 0.02032 0.01317 -0.0830 0.0038 1.0000 6.750 1.0438 0.02169 0.01462 -0.0793 0.0038 1.0000 7.000 1.0525 0.02310 0.01608 -0.0763 0.0038 1.0000 7.250 1.0657 0.02421 0.01724 -0.0743 0.0035 1.0000 7.500 1.0787 0.02540 0.01846 -0.0726 0.0031 1.0000 7.750 1.1052 0.02715 0.02031 -0.0720 0.0026 1.0000 8.000 1.1493 0.02963 0.02289 -0.0741 0.0024 1.0000 8.250 1.2028 0.03320 0.02666 -0.0781 0.0023 1.0000 8.500 1.2334 0.03642 0.03015 -0.0782 0.0024 1.0000 8.750 1.2526 0.03931 0.03326 -0.0769 0.0026 1.0000 9.000 1.2654 0.04232 0.03647 -0.0751 0.0028 1.0000 12.500 1.1566 0.08617 0.08299 -0.0387 0.0033 1.0000 16.000 0.8007 0.17457 0.17287 -0.0866 0.0050 1.0000 16.250 0.8011 0.17973 0.17804 -0.0890 0.0050 1.0000