XFOIL Version 6.94 Calculated polar for: agora02 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.022 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.1543 0.04309 0.03233 -0.0104 0.9988 0.5162 -2.750 -0.1285 0.04222 0.03144 -0.0108 0.9988 0.5350 -2.500 -0.1020 0.04138 0.03065 -0.0113 0.9988 0.5510 -2.250 -0.0721 0.04078 0.02997 -0.0128 0.9988 0.5682 -2.000 -0.0457 0.04014 0.02955 -0.0130 0.9988 0.5894 -1.750 -0.0194 0.03960 0.02930 -0.0131 0.9988 0.6231 -1.500 0.0044 0.03891 0.02922 -0.0123 0.9988 0.6724 -1.250 0.0250 0.03781 0.02918 -0.0100 0.9988 0.7817 -1.000 0.0668 0.03754 0.02853 -0.0161 0.9988 1.0012 -0.500 0.2610 0.04818 0.03807 -0.0261 0.6827 1.0012 -0.250 0.2628 0.05398 0.04359 -0.0197 0.5957 1.0012 0.000 0.2900 0.05627 0.04576 -0.0212 0.5363 1.0012 0.250 0.3393 0.05642 0.04589 -0.0305 0.4928 1.0012 0.500 0.3662 0.05644 0.04567 -0.0277 0.4742 1.0012 0.750 0.3936 0.05707 0.04613 -0.0261 0.4587 1.0012 1.000 0.4247 0.05984 0.04880 -0.0293 0.4461 1.0012 1.250 0.4519 0.06157 0.05041 -0.0290 0.4371 1.0012 1.500 0.4793 0.06446 0.05318 -0.0307 0.4311 1.0012 1.750 0.5047 0.06990 0.05865 -0.0379 0.4235 1.0012 2.000 0.5285 0.07116 0.05980 -0.0364 0.4082 1.0012 2.250 0.5505 0.07328 0.06178 -0.0358 0.3955 1.0012 2.500 0.5670 0.07938 0.06790 -0.0422 0.3936 1.0012 2.750 0.5814 0.08498 0.07348 -0.0467 0.3961 1.0012 3.000 0.5963 0.09009 0.07855 -0.0496 0.3992 1.0012 3.250 0.5700 0.09897 0.08750 -0.0600 0.4191 1.0012 3.500 0.5779 0.10414 0.09260 -0.0624 0.4292 1.0012 3.750 0.5524 0.11012 0.09855 -0.0657 0.4529 1.0012 4.000 0.5347 0.11688 0.10530 -0.0695 0.4887 1.0012 4.250 0.5155 0.12582 0.11426 -0.0764 0.5583 1.0012 4.500 0.5138 0.12851 0.11690 -0.0775 0.5671 1.0012 4.750 0.5274 0.13295 0.12128 -0.0792 0.5712 1.0012 5.000 0.5280 0.13564 0.12390 -0.0800 0.5772 1.0012