XFOIL Version 6.94 Calculated polar for: BOEING 103 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 ------- -------- --------- --------- -------- ------- ------- 1.000 0.5751 0.00755 0.00279 -0.0879 0.6396 0.9929 1.500 0.6434 0.00772 0.00281 -0.0907 0.6154 1.0000 2.000 0.6884 0.00794 0.00285 -0.0885 0.5883 1.0000 2.500 0.7332 0.00817 0.00295 -0.0863 0.5594 1.0000 3.000 0.7770 0.00846 0.00311 -0.0838 0.5238 1.0000 3.500 0.8076 0.00930 0.00334 -0.0788 0.4152 1.0000 4.000 0.8362 0.01064 0.00393 -0.0740 0.2973 1.0000 4.500 0.8783 0.01141 0.00440 -0.0716 0.2433 1.0000 5.000 0.9059 0.01327 0.00541 -0.0671 0.1119 1.0000 5.500 0.9449 0.01438 0.00622 -0.0644 0.0727 1.0000 6.000 0.9805 0.01571 0.00725 -0.0613 0.0262 1.0000 6.500 1.0208 0.01666 0.00815 -0.0588 0.0041 1.0000 7.000 1.0619 0.01740 0.00895 -0.0564 0.0041 1.0000 7.500 1.1001 0.01825 0.00989 -0.0536 0.0042 1.0000 8.000 1.1369 0.01922 0.01096 -0.0507 0.0044 1.0000 8.500 1.1722 0.02029 0.01214 -0.0478 0.0047 1.0000 9.000 1.2046 0.02158 0.01357 -0.0447 0.0051 1.0000 9.500 1.2324 0.02324 0.01540 -0.0413 0.0054 1.0000 10.000 1.2629 0.02476 0.01705 -0.0385 0.0060 1.0000 10.500 1.2880 0.02676 0.01923 -0.0355 0.0066 1.0000 11.000 1.3100 0.02912 0.02176 -0.0326 0.0072 1.0000 11.500 1.3342 0.03143 0.02422 -0.0302 0.0081 1.0000 12.000 1.3512 0.03448 0.02744 -0.0278 0.0088 1.0000 12.500 1.3715 0.03740 0.03052 -0.0260 0.0100 1.0000 13.000 1.3953 0.04012 0.03339 -0.0247 0.0112 1.0000 13.500 1.4023 0.04472 0.03818 -0.0234 0.0122 1.0000 14.000 1.4336 0.04693 0.04053 -0.0227 0.0139 1.0000 14.500 1.4430 0.05167 0.04550 -0.0221 0.0150 1.0000 15.000 1.4464 0.05738 0.05145 -0.0221 0.0159 1.0000 15.500 1.4328 0.06557 0.05990 -0.0230 0.0163 1.0000 16.000 1.4111 0.07530 0.06989 -0.0249 0.0166 1.0000 16.500 1.3879 0.08576 0.08060 -0.0277 0.0167 1.0000 17.000 1.3670 0.09622 0.09131 -0.0309 0.0170 1.0000 17.500 1.3482 0.10658 0.10189 -0.0343 0.0172 1.0000 18.000 1.3333 0.11627 0.11177 -0.0376 0.0174 1.0000 18.500 1.3261 0.12454 0.12019 -0.0403 0.0176 1.0000 19.000 1.3258 0.13144 0.12722 -0.0424 0.0176 1.0000 19.500 1.3289 0.13764 0.13356 -0.0443 0.0175 1.0000 20.000 1.3318 0.14404 0.14009 -0.0466 0.0173 1.0000 20.500 1.3385 0.14930 0.14544 -0.0482 0.0169 1.0000 21.500 1.3180 0.16853 0.16517 -0.0592 0.0166 1.0000