XFOIL Version 6.96 Calculated polar for: FX LIII-142 K 25 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 ------ -------- --------- --------- -------- -------- -------- -18.750 -1.0826 0.12303 0.11952 0.0109 1.0000 0.0086 -18.500 -1.1036 0.11499 0.11135 0.0067 1.0000 0.0087 -18.250 -1.1225 0.10753 0.10376 0.0029 1.0000 0.0087 -18.000 -1.1432 0.09997 0.09607 -0.0008 1.0000 0.0087 -17.750 -1.1633 0.09265 0.08861 -0.0044 1.0000 0.0089 -17.500 -1.1757 0.08683 0.08267 -0.0072 1.0000 0.0087 -17.250 -1.1923 0.08050 0.07620 -0.0101 1.0000 0.0090 -17.000 -1.2025 0.07530 0.07088 -0.0125 1.0000 0.0088 -16.750 -1.2153 0.06990 0.06535 -0.0148 1.0000 0.0094 -16.500 -1.2250 0.06509 0.06041 -0.0169 1.0000 0.0096 -16.250 -1.2314 0.06091 0.05611 -0.0185 1.0000 0.0098 -16.000 -1.2362 0.05714 0.05223 -0.0198 1.0000 0.0100 -15.750 -1.2477 0.05269 0.04767 -0.0212 1.0000 0.0105 -15.500 -1.2530 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0.6584 3.500 0.3911 0.00860 0.00264 0.0018 0.4773 0.6650 3.750 0.4187 0.00861 0.00270 0.0020 0.4637 0.6715 4.000 0.4461 0.00867 0.00274 0.0022 0.4462 0.6791 4.500 0.4990 0.00893 0.00293 0.0029 0.4027 0.6953 4.750 0.5246 0.00910 0.00307 0.0033 0.3782 0.7038 5.000 0.5502 0.00931 0.00323 0.0037 0.3482 0.7134 5.250 0.5745 0.00962 0.00342 0.0043 0.3075 0.7240 5.500 0.5972 0.01003 0.00368 0.0051 0.2596 0.7353 5.750 0.6196 0.01042 0.00399 0.0059 0.2188 0.7486 6.000 0.6411 0.01081 0.00434 0.0069 0.1824 0.7653 6.250 0.6630 0.01118 0.00472 0.0080 0.1614 0.8012 6.500 0.6857 0.01146 0.00504 0.0089 0.1460 0.8277 6.750 0.7084 0.01165 0.00535 0.0099 0.1313 0.8639 7.000 0.7401 0.01195 0.00575 0.0089 0.1132 0.9283 7.250 0.7986 0.01266 0.00636 0.0018 0.0848 0.9749 7.500 0.8300 0.01336 0.00695 0.0005 0.0736 1.0000 7.750 0.8525 0.01373 0.00736 0.0014 0.0702 1.0000 8.000 0.8740 0.01418 0.00782 0.0023 0.0669 1.0000 8.250 0.8942 0.01475 0.00838 0.0035 0.0637 1.0000 8.500 0.9112 0.01556 0.00923 0.0050 0.0607 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0.0252 0.0306 1.0000 13.750 1.2209 0.03325 0.02791 0.0250 0.0291 1.0000 14.000 1.2421 0.03380 0.02853 0.0248 0.0263 1.0000 14.250 1.2544 0.03522 0.02993 0.0245 0.0241 1.0000 14.500 1.2657 0.03674 0.03144 0.0243 0.0197 1.0000 14.750 1.2688 0.03912 0.03383 0.0240 0.0168 1.0000 15.000 1.2644 0.04236 0.03709 0.0235 0.0137 1.0000 15.250 1.2627 0.04535 0.04017 0.0229 0.0123 1.0000 15.500 1.2575 0.04879 0.04368 0.0221 0.0110 1.0000 15.750 1.2495 0.05267 0.04765 0.0210 0.0105 1.0000 16.000 1.2392 0.05696 0.05205 0.0196 0.0099 1.0000 16.250 1.2353 0.06062 0.05582 0.0183 0.0097 1.0000 16.500 1.2263 0.06516 0.06049 0.0165 0.0094 1.0000 16.750 1.2162 0.07005 0.06551 0.0144 0.0093 1.0000 17.000 1.2073 0.07488 0.07045 0.0123 0.0091 1.0000 17.250 1.1936 0.08063 0.07634 0.0097 0.0090 1.0000 17.500 1.1809 0.08635 0.08218 0.0070 0.0090 1.0000 17.750 1.1637 0.09296 0.08893 0.0039 0.0088 1.0000 18.000 1.1473 0.09964 0.09573 0.0006 0.0088 1.0000 18.250 1.1273 0.10708 0.10331 -0.0031 0.0087 1.0000 18.500 1.1068 0.11482 0.11118 -0.0070 0.0087 1.0000