XFOIL Version 6.94 Calculated polar for: Agora0 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 -0.1544 0.04443 0.03354 -0.0130 0.9992 0.5119 -2.750 -0.1273 0.04349 0.03256 -0.0138 0.9992 0.5288 -2.500 -0.0994 0.04278 0.03180 -0.0147 0.9992 0.5496 -2.250 -0.0733 0.04203 0.03120 -0.0147 0.9992 0.5688 -2.000 -0.0442 0.04144 0.03070 -0.0157 0.9992 0.5893 -1.750 -0.0160 0.04090 0.03040 -0.0163 0.9992 0.6175 -1.500 0.0102 0.04026 0.03024 -0.0160 0.9992 0.6605 -1.250 0.0313 0.03920 0.03016 -0.0139 0.9992 0.7557 -1.000 0.0675 0.03857 0.02937 -0.0182 0.9992 1.0008 -0.750 0.1020 0.03950 0.02963 -0.0207 0.9992 1.0008 -0.250 0.3087 0.04785 0.03732 -0.0302 0.6610 1.0008 0.000 0.3244 0.05230 0.04158 -0.0277 0.5894 1.0008 0.250 0.3529 0.05450 0.04369 -0.0292 0.5410 1.0008 0.500 0.3944 0.05551 0.04474 -0.0370 0.4998 1.0008 0.750 0.4228 0.05603 0.04508 -0.0357 0.4802 1.0008 1.000 0.4491 0.05608 0.04491 -0.0328 0.4627 1.0008 1.250 0.4746 0.05587 0.04449 -0.0293 0.4467 1.0008 1.500 0.5013 0.05686 0.04533 -0.0280 0.4345 1.0008 1.750 0.5293 0.05914 0.04754 -0.0288 0.4250 1.0008 2.000 0.5563 0.06176 0.05009 -0.0301 0.4150 1.0008 2.250 0.5821 0.06524 0.05353 -0.0325 0.4063 1.0008 2.500 0.6043 0.07166 0.06002 -0.0393 0.4040 1.0008 2.750 0.6186 0.07909 0.06751 -0.0470 0.4075 1.0008 3.000 0.6285 0.08556 0.07401 -0.0526 0.4111 1.0008 3.250 0.6294 0.09245 0.08090 -0.0583 0.4178 1.0008 3.500 0.6231 0.09906 0.08749 -0.0631 0.4276 1.0008 3.750 0.6137 0.10513 0.09353 -0.0667 0.4403 1.0008 4.000 0.5766 0.11174 0.10010 -0.0696 0.4668 1.0008 4.250 0.5505 0.11860 0.10694 -0.0735 0.5055 1.0008 4.750 0.5428 0.13042 0.11874 -0.0826 0.5944 1.0008 5.000 0.5455 0.13348 0.12174 -0.0836 0.6012 1.0008