XFOIL Version 6.94 Calculated polar for: agora10 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.012 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.0054 0.04336 0.03216 0.0039 0.9988 1.0012 -2.750 -0.0166 0.04161 0.03050 0.0042 0.9988 1.0012 -2.500 -0.0388 0.04005 0.02888 0.0057 0.9988 1.0012 -2.250 -0.0322 0.03962 0.02792 0.0028 0.9988 1.0012 -2.000 -0.0087 0.04013 0.02774 -0.0012 0.9988 1.0012 -1.750 0.0152 0.04100 0.02800 -0.0038 0.9988 1.0012 -1.500 0.0373 0.04205 0.02854 -0.0055 0.9988 1.0012 -1.250 0.0580 0.04322 0.02932 -0.0068 0.9988 1.0012 -1.000 0.0774 0.04456 0.03035 -0.0081 0.9988 1.0012 -0.750 0.0953 0.04610 0.03167 -0.0093 0.9988 1.0012 -0.500 0.1113 0.04796 0.03340 -0.0109 0.9988 1.0012 -0.250 0.1243 0.05032 0.03567 -0.0130 0.9988 1.0012 0.000 0.1341 0.05324 0.03852 -0.0157 0.9988 1.0012 0.250 0.1426 0.05641 0.04159 -0.0185 0.9988 1.0012 0.500 0.1516 0.05957 0.04463 -0.0211 0.9988 1.0012 0.750 0.2222 0.06484 0.04969 -0.0369 0.9618 1.0012 1.000 0.2669 0.06927 0.05397 -0.0458 0.9356 1.0012 1.250 0.2763 0.07251 0.05708 -0.0475 0.9350 1.0012 1.500 0.2625 0.07457 0.05905 -0.0445 0.9495 1.0012 1.750 0.2640 0.07757 0.06190 -0.0445 0.9615 1.0012 2.000 0.2426 0.07906 0.06327 -0.0393 0.9833 1.0012 2.250 0.2295 0.08087 0.06496 -0.0356 0.9988 1.0012 2.500 0.2413 0.08391 0.06788 -0.0373 0.9988 1.0012 2.750 0.2532 0.08697 0.07083 -0.0389 0.9988 1.0012 3.000 0.2651 0.09006 0.07379 -0.0406 0.9988 1.0012 3.250 0.2771 0.09316 0.07679 -0.0422 0.9988 1.0012 3.500 0.2890 0.09629 0.07981 -0.0438 0.9988 1.0012 3.750 0.3009 0.09944 0.08285 -0.0453 0.9988 1.0012 4.000 0.3128 0.10260 0.08591 -0.0469 0.9988 1.0012 4.250 0.3248 0.10578 0.08900 -0.0485 0.9988 1.0012 4.500 0.3367 0.10898 0.09211 -0.0500 0.9988 1.0012 4.750 0.3486 0.11220 0.09525 -0.0515 0.9988 1.0012 5.000 0.3606 0.11544 0.09840 -0.0531 0.9988 1.0012