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 = 0.022 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.0240 0.03322 0.02008 -0.0048 0.9989 0.6080 -2.750 -0.0042 0.03268 0.02008 -0.0021 0.9989 0.6680 -2.500 0.0142 0.03197 0.02000 0.0013 0.9989 0.7440 -2.250 0.0429 0.03100 0.01964 0.0026 0.9989 0.8925 -2.000 0.0749 0.03051 0.01901 0.0005 0.9989 1.0011 -1.750 0.1045 0.03081 0.01900 -0.0009 0.9989 1.0011 -1.500 0.1323 0.03121 0.01916 -0.0017 0.9989 1.0011 -1.250 0.1588 0.03168 0.01948 -0.0021 0.9989 1.0011 -1.000 0.1845 0.03222 0.01996 -0.0023 0.9989 1.0011 -0.750 0.2094 0.03284 0.02060 -0.0024 0.9989 1.0011 -0.500 0.2338 0.03356 0.02144 -0.0026 0.9989 1.0011 -0.250 0.2573 0.03449 0.02262 -0.0028 0.9989 1.0011 0.000 0.2763 0.03635 0.02504 -0.0038 0.9989 1.0011 0.250 0.4372 0.03729 0.02442 -0.0213 0.5922 1.0011 0.500 0.4650 0.03922 0.02580 -0.0202 0.5471 1.0011 0.750 0.4938 0.04114 0.02738 -0.0198 0.5112 1.0011 1.000 0.5235 0.04311 0.02919 -0.0201 0.4801 1.0011 1.250 0.5545 0.04531 0.03133 -0.0208 0.4591 1.0011 1.500 0.5846 0.04759 0.03359 -0.0215 0.4392 1.0011 1.750 0.6144 0.05004 0.03611 -0.0226 0.4215 1.0011 2.000 0.6455 0.05301 0.03927 -0.0244 0.4118 1.0011 2.250 0.6762 0.05655 0.04314 -0.0274 0.4056 1.0011 2.500 0.7053 0.05997 0.04670 -0.0291 0.3987 1.0011 2.750 0.7312 0.06433 0.05140 -0.0326 0.3918 1.0011 3.000 0.7540 0.06883 0.05616 -0.0357 0.3862 1.0011 3.250 0.7725 0.07460 0.06224 -0.0402 0.3894 1.0011 3.500 0.7886 0.08047 0.06830 -0.0440 0.3950 1.0011 3.750 0.7839 0.08865 0.07677 -0.0510 0.4104 1.0011 4.000 0.7715 0.09666 0.08491 -0.0569 0.4295 1.0011 4.250 0.7401 0.10521 0.09350 -0.0631 0.4614 1.0011 4.500 0.7086 0.11329 0.10154 -0.0680 0.5078 1.0011