XFOIL Version 6.94 Calculated polar for: WORTMANN FX 79-K-144/17 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.080 Re = 0.300 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- 0.000 0.2107 0.01801 0.01302 -0.0444 0.7562 0.9118 0.500 0.2648 0.01762 0.01253 -0.0430 0.7520 0.9224 1.000 0.3120 0.01750 0.01243 -0.0417 0.7405 0.9365 1.500 0.3683 0.01722 0.01212 -0.0422 0.7327 0.9431 2.000 0.4185 0.01673 0.01160 -0.0413 0.7275 0.9486 2.500 0.4824 0.01609 0.01094 -0.0429 0.7241 0.9515 3.000 0.5365 0.01592 0.01088 -0.0433 0.7140 0.9564 4.000 0.6516 0.01419 0.00922 -0.0439 0.6972 0.9633 4.500 0.7126 0.01307 0.00815 -0.0447 0.6816 0.9652 5.000 0.7786 0.01143 0.00639 -0.0456 0.6444 0.9676 5.500 0.8280 0.01121 0.00587 -0.0442 0.5948 0.9725 6.000 0.8722 0.01154 0.00602 -0.0428 0.5371 0.9754 6.500 0.9081 0.01231 0.00648 -0.0403 0.4661 0.9787 7.000 0.9223 0.01406 0.00765 -0.0352 0.3456 0.9848 7.500 0.9272 0.01736 0.01027 -0.0316 0.2003 0.9955 8.000 0.9364 0.02030 0.01272 -0.0282 0.1006 1.0000 8.500 0.9514 0.02350 0.01551 -0.0258 0.0293 1.0000 9.000 0.9793 0.02595 0.01811 -0.0245 0.0203 1.0000 9.500 0.9975 0.02924 0.02164 -0.0225 0.0176 1.0000 10.000 1.0130 0.03276 0.02535 -0.0202 0.0169 1.0000 10.500 1.0338 0.03595 0.02877 -0.0183 0.0167 1.0000 11.000 1.0598 0.03903 0.03202 -0.0162 0.0166 1.0000 11.500 1.0974 0.04194 0.03515 -0.0144 0.0167 1.0000 12.500 1.1768 0.04953 0.04334 -0.0120 0.0176 1.0000 13.000 1.1950 0.05345 0.04761 -0.0105 0.0183 1.0000 13.500 1.1807 0.06395 0.05924 -0.0064 0.0223 1.0000 14.000 1.1439 0.07492 0.07077 -0.0047 0.0245 1.0000