XFOIL Version 6.94 Calculated polar for: manu02 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 -0.0291 0.09143 0.07175 -0.0177 0.9999 1.0001 -2.750 -0.0284 0.08942 0.07001 -0.0168 0.9999 1.0001 -2.500 -0.0285 0.08741 0.06830 -0.0158 0.9999 1.0001 -2.250 -0.0295 0.08537 0.06655 -0.0146 0.9999 1.0001 -2.000 -0.0316 0.08332 0.06483 -0.0133 0.9999 1.0001 -1.750 -0.0351 0.08121 0.06301 -0.0118 0.9999 1.0001 -1.500 -0.0400 0.07907 0.06121 -0.0101 0.9999 1.0001 -1.250 -0.0456 0.07689 0.05934 -0.0084 0.9999 1.0001 -1.000 -0.0488 0.07487 0.05754 -0.0072 0.9999 1.0001 -0.750 -0.0463 0.07316 0.05589 -0.0071 0.9999 1.0001 -0.500 -0.0351 0.07197 0.05460 -0.0085 0.9999 1.0001 -0.250 -0.0133 0.07143 0.05381 -0.0117 0.9999 1.0001 0.000 0.0171 0.07156 0.05362 -0.0165 0.9999 1.0001 0.250 0.0520 0.07236 0.05407 -0.0221 0.9999 1.0001 0.500 0.0852 0.07390 0.05529 -0.0278 0.9999 1.0001 0.750 0.1070 0.07633 0.05755 -0.0322 0.9999 1.0001 1.000 0.1118 0.07980 0.06088 -0.0350 0.9999 1.0001 1.250 0.1104 0.08372 0.06449 -0.0373 0.9999 1.0001 1.500 0.1143 0.08749 0.06776 -0.0399 0.9999 1.0001 1.750 0.1235 0.09099 0.07066 -0.0428 0.9999 1.0001 2.000 0.1353 0.09435 0.07338 -0.0456 0.9999 1.0001 2.250 0.1485 0.09758 0.07593 -0.0482 0.9999 1.0001 2.500 0.1620 0.10068 0.07834 -0.0505 0.9999 1.0001 2.750 0.1756 0.10367 0.08067 -0.0526 0.9999 1.0001 3.000 0.1891 0.10657 0.08294 -0.0543 0.9999 1.0001 3.250 0.2022 0.10943 0.08519 -0.0559 0.9999 1.0001 3.500 0.2153 0.11220 0.08738 -0.0573 0.9999 1.0001 3.750 0.2283 0.11493 0.08956 -0.0585 0.9999 1.0001 4.000 0.2411 0.11763 0.09173 -0.0597 0.9999 1.0001 4.250 0.2538 0.12029 0.09389 -0.0607 0.9999 1.0001 4.500 0.2664 0.12293 0.09605 -0.0617 0.9999 1.0001 4.750 0.2789 0.12555 0.09821 -0.0626 0.9999 1.0001 5.000 0.2915 0.12814 0.10037 -0.0635 0.9999 1.0001