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 = 0.012 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.0240 0.04714 0.03436 -0.0006 0.9988 1.0012 -2.750 -0.0444 0.04546 0.03277 0.0026 0.9988 1.0012 -2.500 -0.0624 0.04366 0.03091 0.0047 0.9988 1.0012 -2.250 -0.0494 0.04282 0.02959 0.0015 0.9988 1.0012 -2.000 -0.0201 0.04292 0.02899 -0.0030 0.9988 1.0012 -1.750 0.0093 0.04339 0.02882 -0.0058 0.9988 1.0012 -1.500 0.0362 0.04400 0.02889 -0.0074 0.9988 1.0012 -1.250 0.0614 0.04467 0.02915 -0.0083 0.9988 1.0012 -1.000 0.0853 0.04539 0.02957 -0.0088 0.9988 1.0012 -0.750 0.1085 0.04618 0.03013 -0.0092 0.9988 1.0012 -0.500 0.1310 0.04705 0.03086 -0.0096 0.9988 1.0012 -0.250 0.1527 0.04803 0.03180 -0.0100 0.9988 1.0012 0.000 0.1733 0.04922 0.03303 -0.0106 0.9988 1.0012 0.250 0.1915 0.05083 0.03481 -0.0116 0.9988 1.0012 0.500 0.2011 0.05380 0.03808 -0.0141 0.9988 1.0012 0.750 0.1966 0.05887 0.04329 -0.0187 0.9988 1.0012 1.250 0.3812 0.07397 0.05828 -0.0593 0.8456 1.0012 1.500 0.4087 0.07865 0.06286 -0.0633 0.8179 1.0012 1.750 0.4358 0.08339 0.06749 -0.0668 0.7945 1.0012 2.000 0.4370 0.08704 0.07103 -0.0671 0.7924 1.0012 2.250 0.4425 0.09089 0.07478 -0.0683 0.7945 1.0012 2.500 0.4496 0.09480 0.07861 -0.0697 0.7982 1.0012 2.750 0.4312 0.09694 0.08064 -0.0677 0.8185 1.0012 3.000 0.4187 0.09928 0.08290 -0.0665 0.8420 1.0012 3.250 0.4052 0.10150 0.08503 -0.0650 0.8734 1.0012 3.500 0.3790 0.10262 0.08606 -0.0608 0.9191 1.0012 3.750 0.3293 0.10185 0.08516 -0.0504 0.9801 1.0012 4.000 0.3158 0.10339 0.08659 -0.0464 0.9988 1.0012 4.250 0.3273 0.10664 0.08976 -0.0480 0.9988 1.0012 4.500 0.3389 0.10991 0.09294 -0.0496 0.9988 1.0012 4.750 0.3506 0.11318 0.09613 -0.0512 0.9988 1.0012 5.000 0.3623 0.11646 0.09934 -0.0528 0.9988 1.0012