XFOIL Version 6.96 Calculated polar for: FX 63-145 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.500 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -10.750 -0.0951 0.08677 0.08248 -0.0815 0.4953 0.0177 -10.500 -0.0890 0.08418 0.07984 -0.0819 0.4846 0.0182 -10.250 -0.0860 0.08070 0.07632 -0.0829 0.4766 0.0186 -10.000 -0.0847 0.07682 0.07242 -0.0842 0.4689 0.0191 -9.750 -0.0845 0.07264 0.06824 -0.0856 0.4630 0.0196 -9.500 -0.0881 0.06734 0.06298 -0.0878 0.4677 0.0205 -9.250 -0.0930 0.06179 0.05747 -0.0903 0.4670 0.0207 -9.000 -0.1093 0.05221 0.04793 -0.0961 0.4705 0.0210 -8.750 -0.1188 0.04673 0.04243 -0.0993 0.4648 0.0207 -8.500 -0.1360 0.04095 0.03663 -0.1033 0.4679 0.0202 -8.250 -0.1587 0.03622 0.03187 -0.1065 0.4736 0.0203 -8.000 -0.1803 0.03320 0.02876 -0.1075 0.4690 0.0202 -7.750 -0.1900 0.02966 0.02501 -0.1093 0.4678 0.0206 -7.500 -0.2026 0.02478 0.01958 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2.750 0.8971 0.01345 0.00472 -0.1237 0.1134 0.5512 3.000 0.9245 0.01358 0.00491 -0.1237 0.1118 0.5715 3.250 0.9522 0.01367 0.00508 -0.1237 0.1057 0.5888 3.500 0.9787 0.01388 0.00530 -0.1235 0.1045 0.6045 3.750 1.0049 0.01410 0.00553 -0.1233 0.1033 0.6207 4.000 1.0310 0.01433 0.00579 -0.1231 0.1017 0.6368 4.250 1.0570 0.01455 0.00603 -0.1229 0.1004 0.6481 4.500 1.0830 0.01477 0.00626 -0.1227 0.0992 0.6541 4.750 1.1088 0.01498 0.00651 -0.1224 0.0983 0.6604 5.250 1.1598 0.01546 0.00704 -0.1218 0.0971 0.6737 5.500 1.1849 0.01572 0.00734 -0.1215 0.0967 0.6807 5.750 1.2098 0.01598 0.00764 -0.1211 0.0964 0.6868 6.000 1.2333 0.01634 0.00804 -0.1205 0.0952 0.6944 6.250 1.2546 0.01688 0.00859 -0.1196 0.0925 0.7020 6.500 1.2770 0.01728 0.00905 -0.1189 0.0908 0.7097 6.750 1.3000 0.01763 0.00945 -0.1183 0.0900 0.7173 7.000 1.3227 0.01796 0.00986 -0.1176 0.0894 0.7254 7.500 1.3594 0.01918 0.01125 -0.1150 0.0843 0.7449 7.750 1.3768 0.01978 0.01191 -0.1136 0.0816 0.7554 8.000 1.4030 0.01979 0.01199 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1.0000 13.000 1.5375 0.05020 0.04363 -0.0843 0.0034 1.0000 13.250 1.5371 0.05335 0.04690 -0.0844 0.0033 1.0000 13.500 1.5363 0.05665 0.05030 -0.0847 0.0032 1.0000 13.750 1.5346 0.06014 0.05391 -0.0851 0.0032 1.0000 14.000 1.5324 0.06376 0.05765 -0.0856 0.0031 1.0000 14.250 1.5296 0.06760 0.06160 -0.0862 0.0030 1.0000 14.500 1.5262 0.07162 0.06575 -0.0870 0.0030 1.0000 14.750 1.5221 0.07589 0.07014 -0.0880 0.0029 1.0000 15.000 1.5166 0.08047 0.07485 -0.0892 0.0029 1.0000 15.250 1.5111 0.08516 0.07967 -0.0906 0.0028 1.0000 15.500 1.5055 0.09001 0.08465 -0.0921 0.0028 1.0000 15.750 1.4986 0.09517 0.08994 -0.0938 0.0028 1.0000 16.000 1.4913 0.10053 0.09544 -0.0957 0.0028 1.0000 16.250 1.4838 0.10601 0.10105 -0.0978 0.0028 1.0000 16.500 1.4762 0.11160 0.10677 -0.1000 0.0027 1.0000 16.750 1.4673 0.11753 0.11283 -0.1026 0.0027 1.0000 17.000 1.4591 0.12343 0.11885 -0.1052 0.0027 1.0000 17.250 1.4503 0.12951 0.12506 -0.1081 0.0027 1.0000 17.500 1.4419 0.13553 0.13121 -0.1110 0.0027 1.0000