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 = 0.008 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.0815 0.04866 0.03184 -0.0019 0.9988 1.0012 -2.750 -0.0665 0.04837 0.03091 -0.0034 0.9988 1.0012 -2.500 -0.0460 0.04849 0.03038 -0.0048 0.9988 1.0012 -2.250 -0.0241 0.04881 0.03011 -0.0058 0.9988 1.0012 -2.000 -0.0021 0.04926 0.03004 -0.0065 0.9988 1.0012 -1.750 0.0198 0.04978 0.03012 -0.0069 0.9988 1.0012 -1.500 0.0415 0.05037 0.03033 -0.0072 0.9988 1.0012 -1.250 0.0630 0.05103 0.03068 -0.0074 0.9988 1.0012 -1.000 0.0843 0.05176 0.03115 -0.0075 0.9988 1.0012 -0.750 0.1054 0.05255 0.03175 -0.0076 0.9988 1.0012 -0.500 0.1262 0.05342 0.03250 -0.0078 0.9988 1.0012 -0.250 0.1465 0.05439 0.03341 -0.0080 0.9988 1.0012 0.000 0.1662 0.05550 0.03453 -0.0084 0.9988 1.0012 0.250 0.1848 0.05681 0.03592 -0.0089 0.9988 1.0012 0.500 0.2016 0.05845 0.03772 -0.0098 0.9988 1.0012 0.750 0.2144 0.06074 0.04023 -0.0114 0.9988 1.0012 1.000 0.2199 0.06414 0.04381 -0.0141 0.9988 1.0012 1.250 0.2210 0.06823 0.04793 -0.0173 0.9988 1.0012 1.500 0.2239 0.07213 0.05177 -0.0202 0.9988 1.0012 1.750 0.2291 0.07576 0.05533 -0.0229 0.9988 1.0012 2.000 0.2361 0.07925 0.05873 -0.0253 0.9988 1.0012 2.250 0.2442 0.08263 0.06201 -0.0275 0.9988 1.0012 2.500 0.2530 0.08592 0.06520 -0.0296 0.9988 1.0012 2.750 0.2623 0.08917 0.06835 -0.0316 0.9988 1.0012 3.000 0.2721 0.09238 0.07147 -0.0334 0.9988 1.0012 3.250 0.2823 0.09558 0.07457 -0.0353 0.9988 1.0012 3.500 0.2927 0.09877 0.07767 -0.0371 0.9988 1.0012 3.750 0.3033 0.10196 0.08076 -0.0388 0.9988 1.0012 4.000 0.3142 0.10514 0.08385 -0.0405 0.9988 1.0012 4.250 0.3252 0.10832 0.08694 -0.0422 0.9988 1.0012 4.500 0.3363 0.11150 0.09004 -0.0439 0.9988 1.0012 4.750 0.3476 0.11468 0.09314 -0.0456 0.9988 1.0012 5.000 0.3590 0.11787 0.09625 -0.0473 0.9988 1.0012