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 = 0.450 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- 0.000 0.2649 0.01362 0.00912 -0.0560 0.7496 0.8765 0.500 0.3120 0.01329 0.00884 -0.0538 0.7415 0.8853 1.000 0.3671 0.01320 0.00874 -0.0536 0.7344 0.8969 1.500 0.4159 0.01274 0.00827 -0.0517 0.7294 0.9037 2.000 0.4734 0.01229 0.00779 -0.0520 0.7256 0.9086 2.500 0.5346 0.01181 0.00727 -0.0533 0.7209 0.9118 3.000 0.5954 0.01133 0.00690 -0.0549 0.7068 0.9148 3.500 0.6548 0.01031 0.00583 -0.0552 0.6933 0.9180 4.000 0.7097 0.01001 0.00558 -0.0553 0.6720 0.9210 4.500 0.7650 0.00961 0.00484 -0.0552 0.6231 0.9227 5.000 0.8129 0.01003 0.00501 -0.0543 0.5530 0.9252 5.500 0.8525 0.01097 0.00549 -0.0523 0.4648 0.9293 6.000 0.8862 0.01246 0.00641 -0.0501 0.3503 0.9320 6.500 0.8949 0.01504 0.00811 -0.0447 0.1840 0.9345 7.000 0.9154 0.01730 0.00995 -0.0417 0.0918 0.9375 7.500 0.9416 0.01958 0.01200 -0.0400 0.0317 0.9404 8.000 0.9736 0.02135 0.01379 -0.0385 0.0172 0.9439 8.500 1.0096 0.02294 0.01552 -0.0375 0.0149 0.9463 9.500 1.0645 0.02699 0.01990 -0.0334 0.0131 0.9519 10.000 1.0825 0.02963 0.02278 -0.0304 0.0129 0.9558 10.500 1.0982 0.03258 0.02593 -0.0273 0.0129 0.9600 11.000 1.1091 0.03577 0.02934 -0.0235 0.0130 0.9651 11.500 1.1252 0.03885 0.03264 -0.0201 0.0132 0.9717 12.000 1.1487 0.04182 0.03586 -0.0168 0.0136 0.9803 12.500 1.1816 0.04537 0.03977 -0.0140 0.0144 1.0000 13.000 1.2030 0.05119 0.04609 -0.0115 0.0153 1.0000 13.500 1.1945 0.05981 0.05528 -0.0088 0.0163 1.0000