XFOIL Version 6.94 Calculated polar for: S2062 8% 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.080 Re = 0.450 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- 0.000 0.2291 0.00556 0.00188 -0.0540 0.8807 0.9384 0.500 0.2973 0.00551 0.00176 -0.0559 0.8561 0.9753 1.000 0.3784 0.00548 0.00164 -0.0611 0.8261 0.9988 1.500 0.4254 0.00559 0.00160 -0.0591 0.7879 1.0000 2.000 0.4729 0.00578 0.00162 -0.0571 0.7389 1.0000 2.500 0.5217 0.00609 0.00173 -0.0554 0.6781 1.0000 3.000 0.5676 0.00673 0.00190 -0.0532 0.5674 1.0000 3.500 0.6137 0.00756 0.00223 -0.0513 0.4451 1.0000 4.000 0.6613 0.00846 0.00267 -0.0500 0.3341 1.0000 4.500 0.7084 0.00953 0.00324 -0.0488 0.2179 1.0000 5.000 0.7524 0.01114 0.00414 -0.0472 0.0747 1.0000 5.500 0.7987 0.01260 0.00537 -0.0455 0.0243 1.0000 6.000 0.8467 0.01382 0.00679 -0.0440 0.0213 1.0000 6.500 0.8917 0.01541 0.00857 -0.0420 0.0198 1.0000 7.000 0.9348 0.01736 0.01076 -0.0397 0.0189 1.0000 7.500 0.9781 0.01951 0.01310 -0.0376 0.0177 1.0000 8.000 1.0204 0.02245 0.01632 -0.0355 0.0170 1.0000 8.500 1.0599 0.02651 0.02078 -0.0330 0.0165 1.0000 9.000 1.0839 0.03370 0.02865 -0.0295 0.0152 1.0000 9.500 1.0996 0.04015 0.03574 -0.0251 0.0153 1.0000 10.000 0.9763 0.04009 0.03720 -0.0105 0.0206 1.0000 10.500 0.9227 0.04927 0.04672 -0.0079 0.0212 1.0000 11.000 0.8716 0.06119 0.05890 -0.0117 0.0210 1.0000 11.500 0.8171 0.07767 0.07563 -0.0217 0.0209 1.0000 12.500 0.6828 0.12606 0.12418 -0.0488 0.0332 1.0000 13.000 0.6797 0.13585 0.13396 -0.0559 0.0289 1.0000 13.500 0.6860 0.14240 0.14053 -0.0594 0.0255 1.0000 14.000 0.6931 0.14882 0.14697 -0.0617 0.0234 1.0000