XFOIL Version 6.96 Calculated polar for: FX 63-147 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.050 e 6 Ncrit = 5.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -10.000 -0.3606 0.10410 0.09727 -0.0491 1.0001 0.0384 -9.750 -0.3674 0.09944 0.09271 -0.0506 1.0001 0.0373 -9.500 -0.3969 0.08962 0.08295 -0.0577 1.0001 0.0346 -9.250 -0.4043 0.08616 0.07956 -0.0578 1.0001 0.0343 -9.000 -0.4168 0.08270 0.07617 -0.0578 1.0001 0.0340 -8.750 -0.4334 0.07958 0.07312 -0.0570 1.0001 0.0337 -8.500 -0.4543 0.07697 0.07056 -0.0551 1.0001 0.0335 -8.250 -0.4781 0.07492 0.06858 -0.0520 1.0001 0.0333 -8.000 -0.4993 0.07264 0.06632 -0.0493 1.0001 0.0331 -7.750 -0.5180 0.07027 0.06394 -0.0465 1.0001 0.0329 -7.500 -0.5341 0.06791 0.06152 -0.0437 1.0001 0.0326 -7.250 -0.5478 0.06547 0.05899 -0.0410 1.0001 0.0324 -7.000 -0.5392 0.06125 0.05446 -0.0432 0.9934 0.0320 -6.750 -0.5204 0.05669 0.04946 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0.9940 3.250 0.7306 0.02735 0.01593 -0.0762 0.6283 0.9973 3.500 0.7467 0.02777 0.01641 -0.0750 0.6193 0.9999 3.750 0.7674 0.02795 0.01659 -0.0740 0.6124 0.9999 4.000 0.7688 0.02860 0.01732 -0.0700 0.6038 0.9999 4.250 0.7880 0.02882 0.01755 -0.0688 0.5971 0.9999 4.500 0.7864 0.02952 0.01834 -0.0644 0.5888 0.9999 4.750 0.8027 0.02981 0.01870 -0.0627 0.5822 0.9999 5.000 0.7994 0.03053 0.01950 -0.0580 0.5742 0.9999 5.250 0.8117 0.03088 0.01991 -0.0556 0.5673 0.9999 5.500 0.8079 0.03159 0.02069 -0.0509 0.5595 0.9999 5.750 0.8155 0.03201 0.02119 -0.0479 0.5523 0.9999 6.000 0.8137 0.03268 0.02194 -0.0436 0.5447 0.9999 6.250 0.8206 0.03323 0.02256 -0.0407 0.5371 0.9999 6.500 0.8285 0.03390 0.02333 -0.0382 0.5298 0.9999 6.750 0.8361 0.03467 0.02423 -0.0358 0.5219 0.9999 7.000 0.8516 0.03536 0.02502 -0.0347 0.5149 0.9999 7.250 0.8598 0.03642 0.02622 -0.0330 0.5066 0.9999 7.500 0.8781 0.03721 0.02715 -0.0324 0.4995 0.9999 7.750 0.8890 0.03840 0.02851 -0.0314 0.4910 0.9999 8.000 0.8980 0.03975 0.03001 -0.0303 0.4819 0.9999 8.250 0.9265 0.03993 0.03035 -0.0305 0.4744 0.9999 8.500 0.9235 0.04220 0.03279 -0.0290 0.4635 0.9999 8.750 0.9641 0.04159 0.03238 -0.0299 0.4579 0.9999 9.000 0.9513 0.04478 0.03574 -0.0283 0.4461 0.9999 9.250 1.0038 0.04313 0.03430 -0.0294 0.4414 0.9999 9.500 0.9884 0.04655 0.03788 -0.0279 0.4288 0.9999 9.750 0.9813 0.04966 0.04115 -0.0272 0.4166 0.9999 10.000 1.0463 0.04606 0.03782 -0.0276 0.4110 0.9999 10.250 1.0312 0.04967 0.04159 -0.0266 0.3978 0.9999 10.500 1.0223 0.05311 0.04520 -0.0263 0.3848 0.9999 10.750 1.0255 0.05534 0.04760 -0.0261 0.3724 0.9999 11.000 1.0491 0.05499 0.04745 -0.0254 0.3607 0.9999 11.250 1.0794 0.05370 0.04639 -0.0245 0.3480 0.9999 11.500 1.0953 0.05412 0.04698 -0.0239 0.3338 0.9999 11.750 1.1003 0.05585 0.04887 -0.0236 0.3190 0.9999 12.000 1.0986 0.05852 0.05170 -0.0236 0.3034 0.9999 12.250 1.0955 0.06141 0.05473 -0.0239 0.2865 0.9999 12.500 1.0920 0.06445 0.05790 -0.0243 0.2682 0.9999 12.750 1.0968 0.06625 0.05978 -0.0244 0.2477 0.9999 13.000 1.0982 0.06870 0.06228 -0.0246 0.2281 0.9999 13.250 1.0945 0.07215 0.06579 -0.0254 0.2087 0.9999 13.500 1.0954 0.07465 0.06819 -0.0256 0.1874 0.9999 13.750 1.0902 0.07846 0.07197 -0.0266 0.1684 0.9999 14.000 1.0843 0.08250 0.07599 -0.0277 0.1506 0.9999 14.250 1.0787 0.08649 0.07989 -0.0288 0.1353 0.9999 14.500 1.0734 0.09057 0.08393 -0.0300 0.1224 0.9999 14.750 1.0699 0.09462 0.08797 -0.0312 0.1114 0.9999 15.000 1.0664 0.09862 0.09192 -0.0325 0.1019 0.9999 15.250 1.0638 0.10266 0.09597 -0.0339 0.0934 0.9999 15.500 1.0627 0.10646 0.09983 -0.0350 0.0863 0.9999