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 = 0.045 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.0432 0.02617 0.01689 -0.0200 0.9989 0.6300 -2.750 -0.0278 0.02431 0.01637 -0.0145 0.9989 0.8437 -2.500 0.0169 0.02350 0.01517 -0.0188 0.9989 1.0011 -2.250 0.0460 0.02375 0.01522 -0.0193 0.9989 1.0011 -2.000 0.0736 0.02406 0.01546 -0.0193 0.9989 1.0011 -1.750 0.1005 0.02444 0.01583 -0.0192 0.9989 1.0011 -1.500 0.1271 0.02488 0.01635 -0.0191 0.9989 1.0011 -1.250 0.1537 0.02537 0.01701 -0.0189 0.9989 1.0011 -0.750 0.2608 0.02696 0.01616 -0.0224 0.4318 1.0011 -0.500 0.2896 0.02820 0.01681 -0.0221 0.3831 1.0011 -0.250 0.3194 0.02938 0.01758 -0.0221 0.3531 1.0011 0.000 0.3506 0.03046 0.01846 -0.0222 0.3305 1.0011 0.250 0.3821 0.03162 0.01956 -0.0223 0.3132 1.0011 0.500 0.4127 0.03314 0.02083 -0.0224 0.2995 1.0011 0.750 0.4446 0.03442 0.02222 -0.0227 0.2881 1.0011 1.000 0.4748 0.03619 0.02381 -0.0228 0.2778 1.0011 1.250 0.5066 0.03786 0.02569 -0.0233 0.2704 1.0011 1.500 0.5380 0.03962 0.02761 -0.0238 0.2630 1.0011 1.750 0.5680 0.04164 0.02960 -0.0241 0.2558 1.0011 2.000 0.5980 0.04437 0.03241 -0.0247 0.2512 1.0011 2.250 0.6294 0.04659 0.03509 -0.0256 0.2467 1.0011 2.500 0.6596 0.04923 0.03810 -0.0265 0.2422 1.0011 2.750 0.6883 0.05200 0.04113 -0.0273 0.2371 1.0011 3.000 0.7156 0.05500 0.04425 -0.0279 0.2326 1.0011 3.250 0.7414 0.05898 0.04816 -0.0284 0.2291 1.0011 3.750 0.7884 0.07114 0.06195 -0.0356 0.2451 1.0011 4.000 0.8128 0.07702 0.06767 -0.0358 0.2520 1.0011 4.250 0.6158 0.11422 0.10645 -0.0853 0.5501 1.0011 4.500 0.6328 0.11761 0.10981 -0.0839 0.5204 1.0011 4.750 0.6446 0.12108 0.11323 -0.0824 0.4942 1.0011 5.000 0.6542 0.12455 0.11665 -0.0809 0.4691 1.0011