XFOIL Version 6.94 Calculated polar for: pp12 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.040 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.2067 0.06878 0.06160 -0.0192 0.9999 0.2941 -2.750 -0.1886 0.06681 0.05965 -0.0203 0.9999 0.3125 -2.500 -0.1743 0.06466 0.05757 -0.0199 0.9999 0.3348 -2.250 -0.1530 0.06362 0.05652 -0.0217 0.9999 0.3717 -2.000 -0.1466 0.06059 0.05362 -0.0183 0.9999 0.3966 -1.750 -0.1380 0.05831 0.05145 -0.0156 0.9999 0.4393 -1.500 -0.1345 0.05584 0.04911 -0.0110 0.9999 0.4852 -1.250 -0.1376 0.05311 0.04653 -0.0043 0.9999 0.5500 -1.000 -0.1359 0.05072 0.04426 0.0010 0.9999 0.6096 -0.750 -0.1347 0.04792 0.04158 0.0066 0.9999 0.6564 -0.500 -0.1317 0.04519 0.03893 0.0119 0.9999 0.7104 -0.250 -0.1184 0.04306 0.03687 0.0140 0.9999 0.7441 0.000 -0.0955 0.04162 0.03543 0.0137 0.9999 0.7754 0.250 -0.0614 0.04000 0.03383 0.0110 0.9999 0.8062 0.500 -0.0138 0.03960 0.03337 0.0044 0.9999 0.8221 0.750 0.0465 0.03997 0.03371 -0.0058 0.9999 0.8174 1.000 0.1136 0.04221 0.03579 -0.0195 0.9999 0.7549 1.250 0.1800 0.04552 0.03884 -0.0346 0.9999 0.6110 1.500 0.2413 0.04786 0.04069 -0.0467 0.9999 0.4523 1.750 0.2931 0.04878 0.04119 -0.0539 0.9999 0.3446 2.000 0.7450 0.03576 0.02312 -0.1018 0.1896 0.2344 2.250 0.7850 0.03678 0.02388 -0.1038 0.1777 0.2378 2.500 0.8320 0.03787 0.02472 -0.1073 0.1707 0.2479 2.750 0.9163 0.03869 0.02554 -0.1168 0.1651 0.2764 3.000 1.0791 0.03950 0.02698 -0.1398 0.1560 1.0001 3.250 1.1316 0.04125 0.02849 -0.1426 0.1570 1.0001 3.500 1.1726 0.04320 0.03054 -0.1438 0.1592 1.0001 3.750 1.2089 0.04569 0.03323 -0.1441 0.1617 1.0001 4.000 1.2405 0.04825 0.03592 -0.1439 0.1650 1.0001 4.250 1.2672 0.05107 0.03898 -0.1430 0.1659 1.0001 4.500 1.2953 0.05480 0.04280 -0.1428 0.1688 1.0001 4.750 1.3164 0.05765 0.04600 -0.1408 0.1698 1.0001 5.000 1.3280 0.05892 0.04820 -0.1368 0.1792 1.0001