XFOIL Version 6.96 Calculated polar for: NREL's S806A 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 ------ -------- --------- --------- -------- -------- -------- -3.500 -0.1227 0.02084 0.01488 -0.0555 0.7681 0.0525 -3.250 -0.0944 0.02056 0.01462 -0.0557 0.7652 0.0572 -3.000 -0.0674 0.01848 0.01245 -0.0562 0.7622 0.0667 -2.750 -0.0388 0.01774 0.01165 -0.0564 0.7594 0.0746 -2.500 -0.0112 0.01659 0.01055 -0.0570 0.7561 0.0905 -2.250 0.0252 0.01390 0.00703 -0.0542 0.7525 0.0283 -2.000 0.0536 0.01294 0.00600 -0.0542 0.7497 0.0227 -1.750 0.0828 0.01260 0.00565 -0.0545 0.7492 0.0208 -1.500 0.1120 0.01224 0.00527 -0.0548 0.7485 0.0200 -1.250 0.1414 0.01175 0.00471 -0.0551 0.7477 0.0198 -1.000 0.1711 0.01146 0.00430 -0.0555 0.7467 0.0217 -0.750 0.1893 0.00868 0.00434 -0.0546 0.7457 0.8126 -0.500 0.2093 0.00886 0.00467 -0.0521 0.7443 0.8770 -0.250 0.2254 0.00896 0.00485 -0.0486 0.7428 0.9143 0.000 0.2472 0.00888 0.00479 -0.0470 0.7411 0.9287 0.250 0.2719 0.00883 0.00474 -0.0462 0.7388 0.9391 0.500 0.2978 0.00873 0.00463 -0.0457 0.7363 0.9465 0.750 0.3250 0.00871 0.00460 -0.0456 0.7339 0.9530 1.000 0.3533 0.00871 0.00453 -0.0459 0.7317 0.9588 1.250 0.3828 0.00868 0.00452 -0.0464 0.7290 0.9642 1.500 0.4132 0.00862 0.00445 -0.0471 0.7259 0.9688 1.750 0.4442 0.00855 0.00436 -0.0478 0.7225 0.9735 2.000 0.4750 0.00844 0.00421 -0.0485 0.7184 0.9787 2.250 0.5083 0.00833 0.00407 -0.0497 0.7136 0.9826 2.500 0.5419 0.00839 0.00409 -0.0512 0.7070 0.9859 2.750 0.5752 0.00859 0.00430 -0.0527 0.6992 0.9898 3.000 0.6090 0.00835 0.00407 -0.0541 0.6966 0.9942 3.250 0.6424 0.00819 0.00393 -0.0554 0.6928 0.9996 3.500 0.6694 0.00806 0.00381 -0.0555 0.6875 1.0000 3.750 0.6977 0.00792 0.00367 -0.0557 0.6803 1.0000 4.000 0.7267 0.00780 0.00355 -0.0561 0.6725 1.0000 4.250 0.7556 0.00773 0.00352 -0.0564 0.6638 1.0000 4.500 0.7845 0.00770 0.00343 -0.0567 0.6477 1.0000 4.750 0.8137 0.00756 0.00327 -0.0571 0.6218 1.0000 5.000 0.8323 0.00872 0.00335 -0.0560 0.4269 1.0000 5.250 0.8574 0.00939 0.00376 -0.0562 0.3692 1.0000 5.500 0.8743 0.01103 0.00462 -0.0556 0.2257 1.0000 5.750 0.8955 0.01199 0.00518 -0.0553 0.1580 1.0000 6.000 0.9195 0.01253 0.00560 -0.0553 0.1297 1.0000 6.250 0.9420 0.01318 0.00606 -0.0549 0.1042 1.0000 6.500 0.9650 0.01372 0.00651 -0.0547 0.0872 1.0000 6.750 0.9883 0.01420 0.00693 -0.0544 0.0741 1.0000 7.000 1.0104 0.01476 0.00742 -0.0539 0.0604 1.0000 7.250 1.0320 0.01531 0.00791 -0.0533 0.0508 1.0000 7.500 1.0450 0.01600 0.00867 -0.0508 0.0365 1.0000 7.750 1.0575 0.01685 0.00932 -0.0486 0.0170 1.0000 8.000 1.0722 0.01759 0.00997 -0.0467 0.0096 1.0000 8.250 1.0888 0.01811 0.01051 -0.0451 0.0079 1.0000 8.500 1.1032 0.01872 0.01114 -0.0432 0.0070 1.0000 8.750 1.1158 0.01942 0.01188 -0.0411 0.0063 1.0000 9.000 1.1286 0.02021 0.01272 -0.0392 0.0060 1.0000 9.250 1.1401 0.02112 0.01370 -0.0373 0.0057 1.0000 9.500 1.1521 0.02208 0.01473 -0.0355 0.0055 1.0000 9.750 1.1632 0.02314 0.01587 -0.0336 0.0053 1.0000 10.000 1.1719 0.02440 0.01722 -0.0316 0.0052 1.0000 10.250 1.1805 0.02569 0.01861 -0.0296 0.0052 1.0000 10.500 1.1984 0.02654 0.01956 -0.0293 0.0052 1.0000 10.750 1.2058 0.02799 0.02111 -0.0273 0.0052 1.0000 11.000 1.2120 0.02959 0.02282 -0.0253 0.0052 1.0000 11.250 1.2144 0.03153 0.02490 -0.0228 0.0052 1.0000 11.500 1.2165 0.03362 0.02714 -0.0202 0.0053 1.0000 11.750 1.2181 0.03587 0.02957 -0.0175 0.0054 1.0000 12.000 1.2110 0.03955 0.03357 -0.0134 0.0058 1.0000 12.250 1.2064 0.04313 0.03743 -0.0104 0.0060 1.0000 12.500 1.1721 0.05109 0.04595 -0.0052 0.0070 1.0000 12.750 1.1427 0.05756 0.05280 -0.0026 0.0076 1.0000 13.000 1.0969 0.06635 0.06199 -0.0012 0.0084 1.0000 13.250 1.0725 0.07244 0.06831 -0.0017 0.0085 1.0000 13.500 1.0463 0.07912 0.07521 -0.0032 0.0087 1.0000 13.750 1.0267 0.08562 0.08188 -0.0056 0.0087 1.0000 14.000 0.9983 0.09475 0.09123 -0.0101 0.0089 1.0000 14.250 0.9840 0.10217 0.09879 -0.0147 0.0087 1.0000 14.500 0.9709 0.11024 0.10700 -0.0200 0.0086 1.0000 14.750 0.9482 0.12237 0.11931 -0.0282 0.0085 1.0000