XFOIL Version 6.96 Calculated polar for: NREL's S804 Airfoil 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.100 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -9.250 -0.2756 0.10803 0.10389 -0.0383 1.0000 0.1580 -9.000 -0.2709 0.10693 0.10290 -0.0351 0.9985 0.1596 -8.750 -0.4679 0.06059 0.05561 -0.0872 0.9601 0.0820 -8.500 -0.4410 0.05532 0.05025 -0.0918 0.9483 0.0812 -8.250 -0.4197 0.04950 0.04409 -0.0975 0.9349 0.0797 -8.000 -0.3959 0.04373 0.03775 -0.1033 0.9221 0.0783 -7.750 -0.3583 0.03892 0.03233 -0.1091 0.9138 0.0788 -7.500 -0.3233 0.03557 0.02832 -0.1129 0.9009 0.0813 -7.250 -0.2788 0.03249 0.02495 -0.1170 0.8920 0.0838 -7.000 -0.2365 0.03029 0.02262 -0.1197 0.8803 0.0867 -6.750 -0.1973 0.02852 0.02043 -0.1220 0.8658 0.0913 -6.500 -0.1580 0.02673 0.01870 -0.1239 0.8519 0.0962 -6.250 -0.1174 0.02531 0.01699 -0.1258 0.8378 0.1022 -6.000 -0.0816 0.02402 0.01580 -0.1269 0.8225 0.1104 -5.750 -0.0534 0.02302 0.01482 -0.1267 0.8048 0.1189 -5.500 -0.0272 0.02219 0.01403 -0.1262 0.7876 0.1320 -5.250 -0.0017 0.02134 0.01327 -0.1258 0.7715 0.1541 -5.000 0.0241 0.02009 0.01233 -0.1262 0.7571 0.2042 -4.750 0.0508 0.01870 0.01168 -0.1274 0.7429 0.3241 -4.500 0.0766 0.01907 0.01228 -0.1265 0.7277 0.4150 -4.250 0.1036 0.01990 0.01301 -0.1252 0.7147 0.4592 -4.000 0.1305 0.02074 0.01368 -0.1239 0.7021 0.4931 -3.750 0.1527 0.02163 0.01456 -0.1213 0.6891 0.5153 -3.500 0.1792 0.02233 0.01508 -0.1196 0.6789 0.5356 -3.250 0.1981 0.02307 0.01587 -0.1164 0.6665 0.5487 -3.000 0.2235 0.02362 0.01625 -0.1145 0.6576 0.5642 -2.750 0.2452 0.02397 0.01655 -0.1129 0.6463 0.5802 -2.500 0.2697 0.02443 0.01687 -0.1109 0.6384 0.5935 -2.250 0.2896 0.02476 0.01720 -0.1087 0.6286 0.6051 -2.000 0.3175 0.02485 0.01708 -0.1087 0.6202 0.6203 -1.750 0.3367 0.02521 0.01743 -0.1058 0.6124 0.6287 -1.500 0.3631 0.02522 0.01733 -0.1059 0.6037 0.6419 -1.250 0.3864 0.02542 0.01739 -0.1040 0.5975 0.6502 -1.000 0.4126 0.02548 0.01740 -0.1045 0.5896 0.6620 -0.750 0.4329 0.02564 0.01755 -0.1024 0.5830 0.6690 -0.500 0.4667 0.02564 0.01730 -0.1042 0.5775 0.6802 -0.250 0.4858 0.02584 0.01756 -0.1023 0.5714 0.6862 0.000 0.5096 0.02599 0.01769 -0.1018 0.5648 0.6940 0.250 0.5414 0.02597 0.01752 -0.1031 0.5594 0.7023 0.500 0.5663 0.02614 0.01759 -0.1023 0.5548 0.7082 0.750 0.5890 0.02642 0.01795 -0.1021 0.5486 0.7153 1.000 0.6174 0.02660 0.01808 -0.1029 0.5434 0.7223 1.250 0.6426 0.02675 0.01817 -0.1023 0.5394 0.7279 1.500 0.6742 0.02698 0.01826 -0.1033 0.5359 0.7346 1.750 0.6964 0.02752 0.01891 -0.1034 0.5309 0.7414 2.000 0.7168 0.02791 0.01937 -0.1024 0.5261 0.7469 2.250 0.7440 0.02819 0.01962 -0.1027 0.5219 0.7532 2.500 0.7762 0.02842 0.01974 -0.1040 0.5184 0.7604 2.750 0.8015 0.02882 0.02010 -0.1036 0.5155 0.7667 3.000 0.8172 0.02977 0.02126 -0.1027 0.5111 0.7738 3.250 0.8374 0.03060 0.02218 -0.1025 0.5072 0.7804 3.500 0.8578 0.03121 0.02285 -0.1017 0.5038 0.7875 3.750 0.8871 0.03166 0.02327 -0.1025 0.5006 0.7965 4.000 0.9123 0.03194 0.02353 -0.1020 0.4980 0.8042 4.250 0.9362 0.03288 0.02449 -0.1023 0.4949 0.8135 4.500 0.9333 0.03483 0.02675 -0.0991 0.4907 0.8211 4.750 0.9395 0.03674 0.02884 -0.0977 0.4869 0.8311 5.000 0.9492 0.03803 0.03023 -0.0959 0.4836 0.8415 5.250 0.9721 0.03853 0.03075 -0.0953 0.4810 0.8538 5.500 1.0041 0.03856 0.03073 -0.0956 0.4788 0.8677 5.750 1.0334 0.03896 0.03110 -0.0958 0.4769 0.8834 6.000 0.5302 0.09126 0.08464 -0.0922 0.5269 0.8582 6.250 0.5516 0.09268 0.08609 -0.0922 0.5212 0.8759 6.500 0.5816 0.09426 0.08772 -0.0924 0.5183 0.9012 7.750 0.6397 0.10648 0.10000 -0.0992 0.4832 1.0000 8.000 0.6790 0.10884 0.10229 -0.1011 0.4793 1.0000 8.250 0.7271 0.11192 0.10532 -0.1031 0.4773 1.0000 8.500 0.6855 0.11432 0.10777 -0.1028 0.4635 1.0000 8.750 0.7236 0.11662 0.11003 -0.1039 0.4597 1.0000 9.000 0.7738 0.11979 0.11315 -0.1052 0.4578 1.0000 9.250 0.7268 0.12207 0.11549 -0.1052 0.4435 1.0000 9.500 0.7675 0.12446 0.11786 -0.1059 0.4401 1.0000