XFOIL Version 6.94 Calculated polar for: agora06 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.018 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.1849 0.04121 0.03090 0.0187 0.9988 0.7639 -2.750 -0.1651 0.04082 0.03055 0.0201 0.9988 0.7849 -2.500 -0.1360 0.04050 0.03010 0.0178 0.9988 0.7953 -2.250 -0.1080 0.04013 0.02974 0.0165 0.9988 0.8116 -2.000 -0.0777 0.03979 0.02949 0.0145 0.9988 0.8396 -1.750 -0.0379 0.03948 0.02938 0.0104 0.9988 0.9009 -1.500 0.0003 0.03895 0.02873 0.0036 0.9988 1.0012 -1.250 0.0434 0.03945 0.02859 -0.0040 0.9988 1.0012 -1.000 0.0796 0.04022 0.02877 -0.0082 0.9988 1.0012 -0.750 0.1093 0.04105 0.02919 -0.0102 0.9988 1.0012 -0.500 0.1355 0.04195 0.02984 -0.0113 0.9988 1.0012 -0.250 0.1595 0.04300 0.03081 -0.0121 0.9988 1.0012 0.000 0.1801 0.04456 0.03250 -0.0134 0.9988 1.0012 0.250 0.4053 0.05663 0.04436 -0.0400 0.6598 1.0012 0.500 0.4215 0.06113 0.04868 -0.0383 0.6146 1.0012 0.750 0.4418 0.06489 0.05231 -0.0390 0.5802 1.0012 1.000 0.4630 0.06789 0.05524 -0.0409 0.5521 1.0012 1.250 0.4845 0.07080 0.05809 -0.0439 0.5326 1.0012 1.500 0.5049 0.07397 0.06120 -0.0467 0.5207 1.0012 1.750 0.5201 0.07719 0.06438 -0.0499 0.5109 1.0012 2.000 0.5338 0.08034 0.06746 -0.0521 0.5014 1.0012 2.250 0.5518 0.08347 0.07049 -0.0533 0.4934 1.0012 2.500 0.5603 0.08782 0.07480 -0.0559 0.4928 1.0012 2.750 0.5672 0.09220 0.07911 -0.0580 0.4933 1.0012 3.000 0.5728 0.09651 0.08336 -0.0598 0.4939 1.0012 3.250 0.5782 0.10079 0.08757 -0.0614 0.4949 1.0012 3.500 0.5646 0.10547 0.09220 -0.0626 0.5034 1.0012 3.750 0.5704 0.11015 0.09681 -0.0641 0.5091 1.0012 4.000 0.5748 0.11582 0.10244 -0.0666 0.5222 1.0012 4.250 0.5366 0.12104 0.10770 -0.0752 0.6429 1.0012 4.750 0.5468 0.12804 0.11460 -0.0776 0.6572 1.0012 5.000 0.5518 0.13141 0.11790 -0.0784 0.6584 1.0012