XFOIL Version 6.94 Calculated polar for: HSNLF(1)-0213 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.080 Re = 0.450 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- 0.000 0.0865 0.01251 0.00799 -0.0052 0.7476 0.8379 0.500 0.1453 0.01237 0.00780 -0.0062 0.7423 0.8405 1.000 0.2039 0.01220 0.00756 -0.0070 0.7373 0.8422 1.500 0.2627 0.01217 0.00751 -0.0081 0.7313 0.8434 2.000 0.3214 0.01175 0.00710 -0.0090 0.7223 0.8442 2.500 0.3796 0.01126 0.00652 -0.0092 0.7134 0.8447 3.000 0.4380 0.01064 0.00597 -0.0097 0.6993 0.8458 3.500 0.4967 0.01013 0.00547 -0.0102 0.6844 0.8465 4.000 0.5556 0.00967 0.00505 -0.0108 0.6652 0.8468 4.500 0.6139 0.00929 0.00470 -0.0114 0.6266 0.8472 5.000 0.6630 0.00995 0.00467 -0.0107 0.4683 0.8476 5.500 0.7038 0.01186 0.00570 -0.0100 0.2831 0.8485 6.000 0.7457 0.01353 0.00670 -0.0095 0.1535 0.8492 6.500 0.7899 0.01488 0.00767 -0.0090 0.0889 0.8495 7.000 0.8343 0.01610 0.00874 -0.0085 0.0643 0.8497 7.500 0.8790 0.01719 0.00983 -0.0079 0.0534 0.8499 8.500 0.9536 0.01996 0.01267 -0.0050 0.0439 0.8511 9.000 0.9912 0.02143 0.01421 -0.0040 0.0407 0.8513 10.000 1.0607 0.02504 0.01801 -0.0012 0.0365 0.8519 10.500 1.0946 0.02715 0.02011 0.0004 0.0345 0.8522 11.000 1.1297 0.02897 0.02216 0.0018 0.0331 0.8526 11.500 1.1634 0.03102 0.02437 0.0032 0.0317 0.8531 12.000 1.1990 0.03335 0.02674 0.0046 0.0304 0.8537 13.500 1.2629 0.04234 0.03658 0.0098 0.0270 0.8564 14.000 1.2803 0.04639 0.04077 0.0115 0.0259 0.8575 14.500 1.2708 0.05167 0.04656 0.0133 0.0254 0.8594 15.000 1.2556 0.05819 0.05354 0.0143 0.0247 0.8612 15.500 1.2372 0.06549 0.06122 0.0140 0.0241 0.8627 16.500 1.2095 0.08127 0.07753 0.0095 0.0230 0.8679 17.000 1.1774 0.09359 0.09021 0.0038 0.0228 0.8716