XFOIL Version 6.96 Calculated polar for: Smoothed ATR airfoil coordinates obtained using 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 1.000 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -5.000 -0.2919 0.00960 0.00349 -0.0356 0.7318 0.0517 -4.500 -0.2417 0.00947 0.00318 -0.0341 0.6799 0.0575 -4.000 -0.1909 0.00943 0.00292 -0.0327 0.6222 0.0628 -3.500 -0.1405 0.00932 0.00266 -0.0313 0.5711 0.0712 -2.000 0.0089 0.00929 0.00221 -0.0272 0.3780 0.1512 -1.500 0.0598 0.00910 0.00212 -0.0260 0.3532 0.2135 -1.000 0.1054 0.00836 0.00197 -0.0242 0.3370 0.4113 -0.500 0.1525 0.00783 0.00193 -0.0223 0.3249 0.5740 0.000 0.2017 0.00760 0.00198 -0.0207 0.3136 0.6830 0.500 0.2524 0.00756 0.00205 -0.0193 0.3030 0.7481 1.000 0.3011 0.00740 0.00219 -0.0172 0.2961 0.8401 1.500 0.3517 0.00756 0.00239 -0.0156 0.2871 0.8938 2.000 0.4034 0.00772 0.00256 -0.0142 0.2809 0.9223 2.500 0.4540 0.00795 0.00277 -0.0125 0.2744 0.9464 3.000 0.5083 0.00822 0.00298 -0.0118 0.2675 0.9600 3.500 0.5668 0.00841 0.00314 -0.0121 0.2621 0.9698 4.000 0.6366 0.00875 0.00340 -0.0150 0.2551 0.9748 4.500 0.7064 0.00907 0.00369 -0.0178 0.2500 0.9817 5.000 0.7816 0.00934 0.00391 -0.0220 0.2442 0.9838 5.500 0.8494 0.00973 0.00423 -0.0247 0.2374 0.9867 6.000 0.9160 0.00994 0.00446 -0.0271 0.2338 0.9903 6.500 0.9797 0.01020 0.00470 -0.0290 0.2288 0.9938 7.000 1.0454 0.01063 0.00507 -0.0315 0.2223 0.9966 7.500 1.1116 0.01085 0.00534 -0.0340 0.2187 0.9993 8.000 1.1587 0.01112 0.00563 -0.0326 0.2147 1.0000 8.500 1.1921 0.01149 0.00598 -0.0285 0.2097 1.0000 9.000 1.2247 0.01183 0.00637 -0.0242 0.2059 1.0000 9.500 1.2612 0.01210 0.00668 -0.0207 0.2023 1.0000 10.000 1.2986 0.01247 0.00707 -0.0175 0.1978 1.0000