XFOIL Version 6.96 Calculated polar for: RAE 5215 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 = 5.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -9.500 -0.5797 0.08896 0.08353 -0.0217 1.0000 0.0311 -9.250 -0.5867 0.08174 0.07634 -0.0270 1.0000 0.0307 -9.000 -0.6022 0.07314 0.06772 -0.0344 1.0000 0.0300 -8.750 -0.6254 0.06556 0.06002 -0.0409 1.0000 0.0293 -8.500 -0.6380 0.05939 0.05360 -0.0445 1.0000 0.0291 -8.250 -0.6411 0.05440 0.04833 -0.0462 1.0000 0.0291 -8.000 -0.6388 0.04979 0.04339 -0.0471 1.0000 0.0292 -7.750 -0.6323 0.04541 0.03860 -0.0475 1.0000 0.0292 -7.500 -0.6213 0.04162 0.03441 -0.0475 1.0000 0.0294 -7.250 -0.6071 0.03830 0.03069 -0.0472 1.0000 0.0297 -7.000 -0.5905 0.03546 0.02742 -0.0467 1.0000 0.0308 -6.750 -0.5724 0.03248 0.02382 -0.0461 1.0000 0.0323 -6.500 -0.5522 0.03104 0.02235 -0.0455 1.0000 0.0333 -6.250 -0.5315 0.02921 0.02028 -0.0447 1.0000 0.0341 -6.000 -0.5101 0.02741 0.01823 -0.0438 1.0000 0.0350 -5.750 -0.4882 0.02576 0.01628 -0.0428 1.0000 0.0363 -5.500 -0.4668 0.02462 0.01507 -0.0420 1.0000 0.0380 -5.250 -0.4451 0.02365 0.01403 -0.0411 1.0000 0.0402 -5.000 -0.4228 0.02256 0.01276 -0.0400 1.0000 0.0421 -4.750 -0.4013 0.02169 0.01198 -0.0392 1.0000 0.0442 -4.500 -0.3790 0.02092 0.01115 -0.0383 1.0000 0.0471 -4.250 -0.3566 0.02008 0.01033 -0.0376 1.0000 0.0492 -4.000 -0.3338 0.01940 0.00969 -0.0369 1.0000 0.0522 -3.750 -0.3104 0.01880 0.00906 -0.0364 1.0000 0.0562 -3.500 -0.2864 0.01822 0.00851 -0.0361 1.0000 0.0605 -3.250 -0.2615 0.01768 0.00796 -0.0359 1.0000 0.0660 -3.000 -0.2337 0.01715 0.00747 -0.0363 0.9992 0.0773 -2.750 -0.1986 0.01644 0.00694 -0.0383 0.9961 0.1092 -2.500 -0.1637 0.01450 0.00645 -0.0415 0.9949 0.3936 -2.250 -0.1335 0.01437 0.00700 -0.0417 0.9906 0.5780 -2.000 -0.1021 0.01475 0.00751 -0.0417 0.9859 0.6491 -1.750 -0.0737 0.01513 0.00797 -0.0410 0.9807 0.6955 -1.500 -0.0436 0.01529 0.00814 -0.0408 0.9758 0.7188 -1.250 -0.0081 0.01530 0.00805 -0.0425 0.9720 0.7284 -1.000 0.0228 0.01528 0.00800 -0.0432 0.9662 0.7347 -0.750 0.0583 0.01529 0.00795 -0.0448 0.9620 0.7422 -0.500 0.0926 0.01529 0.00793 -0.0462 0.9576 0.7487 -0.250 0.1267 0.01525 0.00789 -0.0475 0.9502 0.7552 0.000 0.1691 0.01514 0.00775 -0.0502 0.9413 0.7624 0.250 0.2112 0.01484 0.00749 -0.0522 0.9275 0.7679 0.500 0.2479 0.01447 0.00711 -0.0528 0.9045 0.7743 0.750 0.2841 0.01404 0.00665 -0.0530 0.8773 0.7808 1.000 0.3109 0.01379 0.00640 -0.0517 0.8477 0.7868 1.250 0.3394 0.01361 0.00619 -0.0509 0.8201 0.7937 1.500 0.3659 0.01349 0.00609 -0.0499 0.7915 0.8003 1.750 0.3920 0.01339 0.00595 -0.0487 0.7522 0.8075 2.000 0.4176 0.01335 0.00576 -0.0473 0.6798 0.8150 2.250 0.4366 0.01395 0.00544 -0.0443 0.4802 0.8224 2.500 0.4547 0.01522 0.00576 -0.0427 0.3017 0.8314 2.750 0.4767 0.01592 0.00608 -0.0418 0.2315 0.8404 3.000 0.5007 0.01642 0.00640 -0.0411 0.2005 0.8509 3.250 0.5247 0.01682 0.00672 -0.0403 0.1823 0.8636 3.500 0.5482 0.01721 0.00707 -0.0394 0.1694 0.8795 3.750 0.5726 0.01751 0.00741 -0.0386 0.1588 0.9017 4.000 0.5998 0.01791 0.00779 -0.0385 0.1497 0.9453 4.250 0.6288 0.01840 0.00827 -0.0391 0.1409 1.0000 4.500 0.6563 0.01906 0.00883 -0.0393 0.1338 1.0000 4.750 0.6839 0.01966 0.00940 -0.0395 0.1266 1.0000 5.000 0.7105 0.02042 0.01005 -0.0396 0.1206 1.0000 5.250 0.7379 0.02104 0.01075 -0.0397 0.1140 1.0000 5.500 0.7641 0.02180 0.01141 -0.0397 0.1085 1.0000 5.750 0.7909 0.02251 0.01221 -0.0396 0.1022 1.0000 6.000 0.8168 0.02319 0.01286 -0.0395 0.0965 1.0000 6.250 0.8428 0.02396 0.01371 -0.0393 0.0908 1.0000 6.500 0.8682 0.02461 0.01436 -0.0391 0.0855 1.0000 6.750 0.8936 0.02547 0.01532 -0.0389 0.0806 1.0000 7.000 0.9185 0.02619 0.01609 -0.0386 0.0761 1.0000 7.250 0.9431 0.02713 0.01711 -0.0383 0.0722 1.0000 7.500 0.9674 0.02801 0.01811 -0.0379 0.0682 1.0000 7.750 0.9909 0.02891 0.01901 -0.0375 0.0655 1.0000 8.000 1.0143 0.03025 0.02061 -0.0369 0.0619 1.0000 8.250 1.0369 0.03111 0.02151 -0.0364 0.0595 1.0000 8.500 1.0586 0.03255 0.02314 -0.0358 0.0570 1.0000 8.750 1.0793 0.03420 0.02504 -0.0350 0.0548 1.0000 9.000 1.0997 0.03557 0.02654 -0.0343 0.0533 1.0000 9.250 1.1199 0.03675 0.02774 -0.0336 0.0521 1.0000 9.500 1.1341 0.03950 0.03099 -0.0323 0.0503 1.0000 9.750 1.1474 0.04190 0.03375 -0.0310 0.0488 1.0000 10.000 1.1612 0.04376 0.03581 -0.0298 0.0475 1.0000 10.250 1.1753 0.04535 0.03752 -0.0287 0.0466 1.0000 10.500 1.1880 0.04717 0.03942 -0.0276 0.0460 1.0000 10.750 1.1798 0.05184 0.04472 -0.0250 0.0454 1.0000 11.000 1.1636 0.05675 0.05014 -0.0224 0.0449 1.0000 11.250 1.1378 0.06159 0.05536 -0.0196 0.0447 1.0000 11.500 1.1048 0.06749 0.06159 -0.0186 0.0447 1.0000 11.750 1.0628 0.07577 0.07017 -0.0213 0.0449 1.0000