XFOIL Version 6.96 Calculated polar for: USA-35B AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 1.000 e 6 Ncrit = 5.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -15.750 -0.4664 0.15126 0.14940 -0.0086 1.0000 0.0131 -15.500 -0.4600 0.14803 0.14617 -0.0100 1.0000 0.0134 -14.000 -0.8629 0.03452 0.03188 -0.0835 0.9996 0.0206 -13.750 -0.8468 0.03088 0.02794 -0.0871 0.9942 0.0208 -13.500 -0.8266 0.02855 0.02538 -0.0890 0.9879 0.0209 -13.250 -0.8035 0.02658 0.02321 -0.0907 0.9828 0.0210 -13.000 -0.7819 0.02485 0.02129 -0.0915 0.9754 0.0211 -12.750 -0.7622 0.02270 0.01891 -0.0922 0.9671 0.0213 -12.500 -0.7393 0.02149 0.01757 -0.0924 0.9583 0.0215 -12.250 -0.7166 0.02058 0.01654 -0.0922 0.9482 0.0216 -12.000 -0.6932 0.01980 0.01566 -0.0919 0.9392 0.0217 -11.750 -0.6694 0.01910 0.01486 -0.0916 0.9303 0.0219 -11.500 -0.6451 0.01844 0.01411 -0.0913 0.9217 0.0220 -11.250 -0.6207 0.01782 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0.2722 1.0000 3.750 1.0274 0.00913 0.00347 -0.0851 0.2559 1.0000 4.000 1.0532 0.00942 0.00367 -0.0848 0.2387 1.0000 4.250 1.0787 0.00974 0.00389 -0.0844 0.2206 1.0000 4.750 1.1276 0.01054 0.00445 -0.0834 0.1790 1.0000 5.000 1.1521 0.01091 0.00475 -0.0828 0.1640 1.0000 5.250 1.1770 0.01124 0.00502 -0.0824 0.1538 1.0000 5.500 1.2018 0.01156 0.00529 -0.0819 0.1453 1.0000 5.750 1.2270 0.01183 0.00555 -0.0815 0.1390 1.0000 6.000 1.2515 0.01215 0.00584 -0.0810 0.1321 1.0000 6.250 1.2760 0.01246 0.00613 -0.0805 0.1250 1.0000 6.500 1.2996 0.01284 0.00645 -0.0798 0.1168 1.0000 6.750 1.3239 0.01314 0.00675 -0.0793 0.1116 1.0000 7.000 1.3467 0.01354 0.00710 -0.0785 0.1037 1.0000 7.250 1.3691 0.01396 0.00748 -0.0777 0.0931 1.0000 7.750 1.4008 0.01572 0.00893 -0.0743 0.0471 1.0000 8.000 1.4211 0.01620 0.00941 -0.0732 0.0430 1.0000 8.250 1.4403 0.01673 0.00993 -0.0719 0.0395 1.0000 8.500 1.4601 0.01717 0.01040 -0.0707 0.0375 1.0000 8.750 1.4777 0.01766 0.01090 -0.0692 0.0356 1.0000 9.000 1.4931 0.01821 0.01146 -0.0673 0.0337 1.0000 9.250 1.5096 0.01872 0.01200 -0.0656 0.0326 1.0000 9.500 1.5255 0.01927 0.01258 -0.0640 0.0315 1.0000 9.750 1.5403 0.01990 0.01323 -0.0622 0.0303 1.0000 10.000 1.5541 0.02061 0.01396 -0.0605 0.0291 1.0000 10.250 1.5678 0.02135 0.01474 -0.0588 0.0282 1.0000 10.500 1.5822 0.02207 0.01551 -0.0573 0.0275 1.0000 10.750 1.5956 0.02288 0.01636 -0.0558 0.0266 1.0000 11.000 1.6079 0.02379 0.01731 -0.0542 0.0257 1.0000 11.250 1.6191 0.02483 0.01838 -0.0527 0.0248 1.0000 11.500 1.6289 0.02600 0.01960 -0.0512 0.0239 1.0000 11.750 1.6406 0.02708 0.02074 -0.0500 0.0234 1.0000 12.000 1.6511 0.02829 0.02201 -0.0488 0.0228 1.0000 12.250 1.6605 0.02964 0.02340 -0.0476 0.0220 1.0000 12.500 1.6687 0.03114 0.02496 -0.0465 0.0212 1.0000 12.750 1.6753 0.03284 0.02670 -0.0454 0.0205 1.0000 13.000 1.6817 0.03461 0.02854 -0.0445 0.0198 1.0000 13.250 1.6882 0.03644 0.03044 -0.0437 0.0193 1.0000 13.500 1.6934 0.03846 0.03252 -0.0431 0.0187 1.0000 13.750 1.6973 0.04067 0.03480 -0.0425 0.0181 1.0000 14.000 1.6995 0.04311 0.03730 -0.0421 0.0175 1.0000 14.250 1.7005 0.04578 0.04003 -0.0417 0.0169 1.0000 14.500 1.7015 0.04850 0.04284 -0.0415 0.0165 1.0000 14.750 1.7017 0.05137 0.04579 -0.0414 0.0161 1.0000 15.000 1.7007 0.05444 0.04894 -0.0415 0.0156 1.0000 15.250 1.6984 0.05772 0.05230 -0.0416 0.0151 1.0000 15.500 1.6947 0.06122 0.05587 -0.0418 0.0147 1.0000 15.750 1.6893 0.06496 0.05970 -0.0422 0.0143 1.0000 16.000 1.6827 0.06894 0.06376 -0.0427 0.0139 1.0000 16.250 1.6766 0.07293 0.06784 -0.0433 0.0136 1.0000 16.500 1.6697 0.07707 0.07208 -0.0440 0.0134 1.0000 16.750 1.6617 0.08142 0.07653 -0.0448 0.0131 1.0000 17.000 1.6528 0.08598 0.08118 -0.0458 0.0129 1.0000 17.250 1.6431 0.09074 0.08604 -0.0470 0.0126 1.0000 17.500 1.6324 0.09569 0.09109 -0.0483 0.0124 1.0000 17.750 1.6210 0.10082 0.09631 -0.0498 0.0122 1.0000 18.000 1.6090 0.10613 0.10172 -0.0515 0.0120 1.0000 18.250 1.5965 0.11160 0.10729 -0.0534 0.0118 1.0000 18.500 1.5833 0.11721 0.11300 -0.0554 0.0116 1.0000 18.750 1.5698 0.12296 0.11884 -0.0577 0.0114 1.0000