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 = 0.008 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.0260 0.04596 0.02465 0.0015 0.9989 1.0011 -2.750 -0.0018 0.04589 0.02421 0.0012 0.9989 1.0011 -2.500 0.0226 0.04591 0.02391 0.0010 0.9989 1.0011 -2.250 0.0471 0.04601 0.02372 0.0008 0.9989 1.0011 -2.000 0.0715 0.04619 0.02367 0.0007 0.9989 1.0011 -1.750 0.0956 0.04644 0.02374 0.0006 0.9989 1.0011 -1.500 0.1195 0.04676 0.02395 0.0006 0.9989 1.0011 -1.250 0.1429 0.04715 0.02427 0.0006 0.9989 1.0011 -1.000 0.1659 0.04762 0.02474 0.0006 0.9989 1.0011 -0.750 0.1885 0.04818 0.02536 0.0006 0.9989 1.0011 -0.500 0.2105 0.04883 0.02615 0.0006 0.9989 1.0011 -0.250 0.2320 0.04961 0.02711 0.0006 0.9989 1.0011 0.000 0.2526 0.05055 0.02831 0.0005 0.9989 1.0011 0.250 0.2717 0.05172 0.02984 0.0003 0.9989 1.0011 0.500 0.2881 0.05333 0.03187 -0.0002 0.9989 1.0011 0.750 0.2989 0.05584 0.03483 -0.0012 0.9989 1.0011 1.000 0.2980 0.06008 0.03938 -0.0034 0.9989 1.0011 1.250 0.2895 0.06552 0.04487 -0.0067 0.9989 1.0011 1.500 0.2852 0.07047 0.04977 -0.0099 0.9989 1.0011 1.750 0.2857 0.07480 0.05404 -0.0128 0.9989 1.0011 2.000 0.2891 0.07875 0.05792 -0.0153 0.9989 1.0011 2.250 0.2947 0.08252 0.06162 -0.0178 0.9989 1.0011 2.500 0.3016 0.08614 0.06517 -0.0200 0.9989 1.0011 2.750 0.3093 0.08967 0.06864 -0.0222 0.9989 1.0011 3.000 0.3176 0.09314 0.07204 -0.0242 0.9989 1.0011 3.250 0.3264 0.09658 0.07541 -0.0262 0.9989 1.0011 3.500 0.3356 0.10000 0.07877 -0.0281 0.9989 1.0011 3.750 0.3452 0.10340 0.08211 -0.0300 0.9989 1.0011 4.000 0.3549 0.10679 0.08544 -0.0319 0.9989 1.0011 4.250 0.3649 0.11018 0.08877 -0.0337 0.9989 1.0011 4.500 0.3751 0.11356 0.09210 -0.0356 0.9989 1.0011 4.750 0.3854 0.11695 0.09544 -0.0374 0.9989 1.0011 5.000 0.3959 0.12033 0.09878 -0.0392 0.9989 1.0011