XFOIL Version 6.96 Calculated polar for: DSMA-523B 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 ------ -------- --------- --------- -------- -------- -------- -9.000 -0.5589 0.10034 0.09466 -0.0156 1.0000 0.1893 -8.750 -0.7755 0.05484 0.04758 -0.0538 1.0000 0.0969 -8.500 -0.7587 0.05064 0.04321 -0.0547 1.0000 0.0960 -8.250 -0.7396 0.04641 0.03871 -0.0562 1.0000 0.0953 -8.000 -0.7170 0.04239 0.03436 -0.0580 1.0000 0.0944 -7.750 -0.6909 0.03865 0.03022 -0.0600 1.0000 0.0935 -7.500 -0.6626 0.03555 0.02669 -0.0617 1.0000 0.0936 -7.250 -0.6329 0.03320 0.02391 -0.0632 1.0000 0.0953 -7.000 -0.6042 0.03112 0.02154 -0.0642 1.0000 0.0978 -6.750 -0.5785 0.02962 0.02003 -0.0643 1.0000 0.1005 -6.500 -0.5516 0.02827 0.01859 -0.0642 1.0000 0.1031 -6.250 -0.5237 0.02713 0.01728 -0.0643 1.0000 0.1070 -6.000 -0.4973 0.02588 0.01601 -0.0642 1.0000 0.1115 -5.750 -0.4708 0.02497 0.01517 -0.0639 1.0000 0.1165 -5.500 -0.4429 0.02420 0.01430 -0.0638 1.0000 0.1223 -5.250 -0.4166 0.02327 0.01356 -0.0636 1.0000 0.1298 -5.000 -0.3884 0.02268 0.01295 -0.0635 1.0000 0.1386 -4.750 -0.3606 0.02193 0.01244 -0.0637 1.0000 0.1501 -4.500 -0.3313 0.02130 0.01199 -0.0640 1.0000 0.1658 -4.250 -0.2997 0.02062 0.01160 -0.0649 1.0000 0.1979 -4.000 -0.2590 0.01909 0.01225 -0.0682 1.0000 0.4803 -3.750 -0.2712 0.02082 0.01441 -0.0559 1.0000 0.5581 -3.500 -0.2370 0.02181 0.01518 -0.0570 1.0000 0.6318 -3.250 -0.2223 0.02282 0.01613 -0.0528 1.0000 0.6538 -3.000 -0.2057 0.02360 0.01686 -0.0493 1.0000 0.6720 -2.750 -0.1899 0.02427 0.01749 -0.0456 1.0000 0.6890 -2.500 -0.1694 0.02476 0.01792 -0.0434 1.0000 0.7052 -2.250 -0.1628 0.02510 0.01829 -0.0374 1.0000 0.7148 -2.000 -0.1469 0.02538 0.01855 -0.0341 1.0000 0.7294 -1.750 -0.1285 0.02562 0.01878 -0.0316 1.0000 0.7450 -1.500 -0.1153 0.02574 0.01891 -0.0277 1.0000 0.7587 -1.250 -0.1048 0.02572 0.01891 -0.0232 1.0000 0.7722 -1.000 -0.0914 0.02569 0.01891 -0.0196 1.0000 0.7878 -0.750 -0.0797 0.02561 0.01885 -0.0155 1.0000 0.8052 -0.500 -0.0683 0.02544 0.01871 -0.0116 1.0000 0.8228 -0.250 -0.0556 0.02521 0.01852 -0.0081 1.0000 0.8396 0.000 -0.0364 0.02505 0.01838 -0.0064 1.0000 0.8531 0.250 -0.0159 0.02489 0.01825 -0.0051 1.0000 0.8639 0.500 0.0021 0.02461 0.01802 -0.0034 1.0000 0.8732 0.750 0.0261 0.02455 0.01801 -0.0032 1.0000 0.8822 1.000 0.0488 0.02448 0.01801 -0.0027 1.0000 0.8910 1.250 0.0690 0.02434 0.01793 -0.0018 1.0000 0.8990 1.500 0.0930 0.02436 0.01805 -0.0018 1.0000 0.9060 1.750 0.1185 0.02450 0.01829 -0.0022 1.0000 0.9114 2.000 0.2150 0.02384 0.01778 -0.0147 0.9639 0.9158 2.250 0.2846 0.02204 0.01617 -0.0211 0.9268 0.9176 2.500 0.3336 0.02016 0.01448 -0.0227 0.8653 0.9185 2.750 0.4398 0.02161 0.01223 -0.0335 0.1837 0.9199 3.000 0.4681 0.02243 0.01286 -0.0339 0.1624 0.9213 3.250 0.4979 0.02335 0.01359 -0.0347 0.1485 0.9223 3.500 0.5300 0.02403 0.01427 -0.0357 0.1381 0.9227 3.750 0.5631 0.02515 0.01520 -0.0371 0.1305 0.9235 4.000 0.5961 0.02611 0.01624 -0.0381 0.1248 0.9248 4.250 0.6285 0.02700 0.01719 -0.0392 0.1184 0.9258 4.500 0.6630 0.02841 0.01854 -0.0408 0.1141 0.9261 4.750 0.6971 0.03036 0.02061 -0.0423 0.1109 0.9264 5.000 0.7273 0.03150 0.02207 -0.0427 0.1073 0.9273 5.250 0.7573 0.03309 0.02392 -0.0432 0.1046 0.9283 5.500 0.7866 0.03508 0.02622 -0.0436 0.1031 0.9289 5.750 0.8146 0.03713 0.02853 -0.0439 0.1014 0.9290 6.000 0.8429 0.03917 0.03063 -0.0447 0.0987 0.9292 6.250 0.8674 0.04225 0.03393 -0.0449 0.0973 0.9294 6.500 0.8792 0.04650 0.03931 -0.0418 0.1026 0.9295 6.750 0.8858 0.05365 0.04739 -0.0390 0.1184 0.9297 7.000 0.8444 0.07323 0.06877 -0.0382 0.2134 0.9295 7.250 0.8298 0.07781 0.07354 -0.0379 0.2022 0.9296 7.500 0.8369 0.08094 0.07675 -0.0378 0.1939 0.9302 7.750 0.8989 0.08800 0.08332 -0.0391 0.1879 0.9315 8.000 0.8057 0.09050 0.08646 -0.0388 0.1817 0.9311 8.250 0.7955 0.09534 0.09137 -0.0414 0.1749 0.9312 8.500 0.9086 0.09994 0.09553 -0.0389 0.1675 0.9323 8.750 0.8208 0.10388 0.09987 -0.0409 0.1659 0.9315 9.000 0.7709 0.11317 0.10918 -0.0548 0.1581 0.9311