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 = 0.005 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.0608 0.05737 0.03310 -0.0068 0.9988 1.0012 -2.750 -0.0436 0.05752 0.03269 -0.0069 0.9988 1.0012 -2.500 -0.0258 0.05779 0.03245 -0.0070 0.9988 1.0012 -2.250 -0.0076 0.05816 0.03234 -0.0069 0.9988 1.0012 -2.000 0.0110 0.05861 0.03236 -0.0068 0.9988 1.0012 -1.750 0.0298 0.05912 0.03250 -0.0067 0.9988 1.0012 -1.500 0.0486 0.05971 0.03276 -0.0066 0.9988 1.0012 -1.250 0.0675 0.06036 0.03313 -0.0065 0.9988 1.0012 -1.000 0.0864 0.06108 0.03361 -0.0064 0.9988 1.0012 -0.750 0.1052 0.06187 0.03422 -0.0063 0.9988 1.0012 -0.500 0.1239 0.06275 0.03496 -0.0063 0.9988 1.0012 -0.250 0.1424 0.06372 0.03585 -0.0064 0.9988 1.0012 0.000 0.1606 0.06480 0.03690 -0.0065 0.9988 1.0012 0.250 0.1783 0.06600 0.03813 -0.0068 0.9988 1.0012 0.500 0.1953 0.06738 0.03959 -0.0072 0.9988 1.0012 0.750 0.2110 0.06898 0.04132 -0.0079 0.9988 1.0012 1.000 0.2249 0.07090 0.04341 -0.0089 0.9988 1.0012 1.250 0.2361 0.07326 0.04595 -0.0103 0.9988 1.0012 1.500 0.2439 0.07617 0.04897 -0.0122 0.9988 1.0012 1.750 0.2491 0.07949 0.05232 -0.0145 0.9988 1.0012 2.000 0.2538 0.08296 0.05574 -0.0169 0.9988 1.0012 2.250 0.2593 0.08637 0.05908 -0.0192 0.9988 1.0012 2.500 0.2657 0.08970 0.06231 -0.0215 0.9988 1.0012 2.750 0.2730 0.09296 0.06547 -0.0236 0.9988 1.0012 3.000 0.2811 0.09618 0.06860 -0.0256 0.9988 1.0012 3.250 0.2898 0.09937 0.07169 -0.0276 0.9988 1.0012 3.500 0.2990 0.10252 0.07474 -0.0295 0.9988 1.0012 3.750 0.3086 0.10565 0.07777 -0.0314 0.9988 1.0012 4.000 0.3184 0.10876 0.08078 -0.0333 0.9988 1.0012 4.250 0.3286 0.11184 0.08378 -0.0351 0.9988 1.0012 4.500 0.3390 0.11492 0.08677 -0.0369 0.9988 1.0012 4.750 0.3496 0.11799 0.08975 -0.0387 0.9988 1.0012 5.000 0.3604 0.12105 0.09273 -0.0405 0.9988 1.0012