XFOIL Version 6.94 Calculated polar for: WBL FX 71-089A AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.018 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.0532 0.03388 0.02142 0.0070 0.9989 0.8478 -2.750 -0.0085 0.03299 0.02076 0.0043 0.9989 1.0011 -2.500 0.0194 0.03274 0.02011 0.0022 0.9989 1.0011 -2.250 0.0484 0.03280 0.01979 0.0006 0.9989 1.0011 -2.000 0.0767 0.03301 0.01967 -0.0004 0.9989 1.0011 -1.750 0.1040 0.03332 0.01972 -0.0010 0.9989 1.0011 -1.500 0.1303 0.03369 0.01991 -0.0014 0.9989 1.0011 -1.250 0.1559 0.03414 0.02024 -0.0016 0.9989 1.0011 -1.000 0.1808 0.03466 0.02072 -0.0017 0.9989 1.0011 -0.750 0.2052 0.03527 0.02136 -0.0018 0.9989 1.0011 -0.500 0.2290 0.03598 0.02218 -0.0019 0.9989 1.0011 -0.250 0.2520 0.03686 0.02328 -0.0020 0.9989 1.0011 0.000 0.2731 0.03814 0.02496 -0.0025 0.9989 1.0011 0.250 0.2748 0.04257 0.03006 -0.0050 0.9989 1.0011 0.500 0.4919 0.04305 0.02943 -0.0325 0.6322 1.0011 0.750 0.5203 0.04528 0.03135 -0.0318 0.5928 1.0011 1.000 0.5485 0.04759 0.03344 -0.0316 0.5606 1.0011 1.250 0.5756 0.04985 0.03550 -0.0310 0.5317 1.0011 1.500 0.6051 0.05257 0.03819 -0.0320 0.5116 1.0011 1.750 0.6340 0.05559 0.04129 -0.0336 0.4939 1.0011 2.000 0.6608 0.05877 0.04456 -0.0352 0.4768 1.0011 2.250 0.6866 0.06262 0.04860 -0.0379 0.4661 1.0011 2.500 0.7103 0.06721 0.05344 -0.0417 0.4622 1.0011 2.750 0.7312 0.07240 0.05889 -0.0462 0.4614 1.0011 3.000 0.7451 0.07788 0.06455 -0.0501 0.4611 1.0011 3.250 0.7525 0.08352 0.07032 -0.0534 0.4614 1.0011 3.500 0.7525 0.08952 0.07640 -0.0566 0.4645 1.0011 3.750 0.7532 0.09538 0.08229 -0.0592 0.4693 1.0011 4.000 0.7335 0.10228 0.08920 -0.0629 0.4855 1.0011 4.250 0.7151 0.10876 0.09566 -0.0659 0.5062 1.0011 4.500 0.6834 0.11509 0.10192 -0.0683 0.5383 1.0011 4.750 0.6725 0.12197 0.10877 -0.0731 0.5856 1.0011 5.000 0.4034 0.11837 0.10488 -0.0500 0.9989 1.0011