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 = 0.012 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.1843 0.05860 0.04574 0.0178 0.9999 0.9101 -2.750 -0.2194 0.05789 0.04527 0.0261 0.9999 0.8976 -2.500 -0.2547 0.05712 0.04476 0.0342 0.9999 0.8851 -2.250 -0.2775 0.05624 0.04403 0.0390 0.9999 0.8673 -2.000 -0.2868 0.05554 0.04343 0.0404 0.9999 0.8455 -1.750 -0.2704 0.05514 0.04298 0.0356 0.9999 0.8137 -1.500 -0.2312 0.05528 0.04293 0.0257 0.9999 0.7767 -1.250 -0.1783 0.05520 0.04261 0.0139 0.9999 0.7409 -1.000 -0.1206 0.05524 0.04241 0.0018 0.9999 0.7078 -0.750 -0.0586 0.05559 0.04244 -0.0108 0.9999 0.6738 -0.500 -0.0001 0.05556 0.04208 -0.0217 0.9999 0.6426 -0.250 0.0524 0.05569 0.04194 -0.0306 0.9999 0.6181 0.000 0.1020 0.05583 0.04182 -0.0382 0.9999 0.6010 0.250 0.1518 0.05580 0.04157 -0.0454 0.9999 0.5909 0.500 0.1983 0.05602 0.04156 -0.0518 0.9999 0.5813 0.750 0.2461 0.05620 0.04160 -0.0580 0.9999 0.5744 1.000 0.2940 0.05656 0.04178 -0.0642 0.9999 0.5702 1.250 0.3394 0.05692 0.04210 -0.0697 0.9999 0.5729 1.500 0.3861 0.05738 0.04251 -0.0753 0.9999 0.5828 1.750 0.4349 0.05774 0.04305 -0.0810 0.9999 0.5979 2.000 0.4793 0.05839 0.04394 -0.0861 0.9999 0.6141 2.250 0.5364 0.05864 0.04492 -0.0933 0.9999 0.6663 2.500 0.5799 0.05831 0.04580 -0.0985 0.9999 0.8086 2.750 0.6306 0.06064 0.04787 -0.1033 0.9999 1.0001 3.000 0.6660 0.06243 0.04954 -0.1058 0.9999 1.0001 3.250 0.6999 0.06407 0.05096 -0.1072 0.9999 1.0001 3.500 0.7292 0.06561 0.05241 -0.1074 0.9999 1.0001 3.750 0.7549 0.06727 0.05430 -0.1071 0.9999 1.0001 4.000 0.7718 0.07004 0.05758 -0.1079 0.9999 1.0001 4.250 1.1189 0.06192 0.04804 -0.1262 0.4405 1.0001 4.500 1.1505 0.06539 0.05132 -0.1266 0.4255 1.0001 4.750 1.1589 0.07006 0.05630 -0.1262 0.4167 1.0001 5.000 1.1678 0.07468 0.06111 -0.1263 0.4131 1.0001