XFOIL Version 6.94 Calculated polar for: ARA-D BL 12% 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.040 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.1738 0.03952 0.03140 -0.0210 1.0000 0.2369 -2.750 -0.1543 0.03772 0.02975 -0.0205 1.0000 0.2451 -2.500 -0.1308 0.03647 0.02834 -0.0204 1.0000 0.2548 -2.250 -0.1060 0.03520 0.02691 -0.0207 1.0000 0.2654 -2.000 -0.0821 0.03435 0.02608 -0.0205 1.0000 0.2794 -1.750 0.1341 0.02693 0.01913 -0.0435 0.8439 0.3779 -1.500 0.1510 0.02648 0.01705 -0.0326 0.4824 0.4139 -1.250 0.1724 0.02683 0.01671 -0.0302 0.3962 0.4660 -1.000 0.1924 0.02602 0.01658 -0.0272 0.3661 0.5898 -0.750 0.2440 0.02577 0.01626 -0.0282 0.3379 1.0000 -0.500 0.2738 0.02683 0.01667 -0.0278 0.3238 1.0000 -0.250 0.3044 0.02780 0.01731 -0.0275 0.3136 1.0000 0.000 0.3346 0.02906 0.01812 -0.0273 0.3061 1.0000 0.250 0.3646 0.03010 0.01911 -0.0271 0.3003 1.0000 0.500 0.3937 0.03129 0.02018 -0.0269 0.2941 1.0000 0.750 0.4219 0.03267 0.02131 -0.0266 0.2877 1.0000 1.000 0.4498 0.03424 0.02278 -0.0264 0.2828 1.0000 1.250 0.4779 0.03570 0.02436 -0.0264 0.2795 1.0000 1.500 0.5056 0.03742 0.02618 -0.0265 0.2777 1.0000 1.750 0.5325 0.03934 0.02822 -0.0266 0.2767 1.0000 2.000 0.5584 0.04148 0.03050 -0.0268 0.2762 1.0000 2.250 0.5830 0.04380 0.03298 -0.0270 0.2754 1.0000 2.500 0.6058 0.04631 0.03566 -0.0273 0.2742 1.0000 2.750 0.6266 0.04914 0.03868 -0.0276 0.2735 1.0000 3.000 0.6447 0.05255 0.04232 -0.0283 0.2754 1.0000 3.250 0.6613 0.05636 0.04630 -0.0291 0.2794 1.0000 3.500 0.6797 0.06010 0.05005 -0.0295 0.2831 1.0000 3.750 0.6569 0.07029 0.06113 -0.0367 0.3115 1.0000 4.500 0.4624 0.11010 0.10177 -0.0776 0.6024 1.0000 5.000 0.4606 0.11308 0.10453 -0.0739 0.5620 1.0000