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 = 0.035 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.2364 0.06881 0.06134 -0.0089 0.9999 0.3825 -2.750 -0.2228 0.06689 0.05947 -0.0088 0.9999 0.4166 -2.500 -0.2162 0.06452 0.05721 -0.0060 0.9999 0.4586 -2.250 -0.2170 0.06126 0.05408 -0.0003 0.9999 0.5046 -2.000 -0.2214 0.05832 0.05130 0.0066 0.9999 0.5674 -1.750 -0.2258 0.05495 0.04808 0.0136 0.9999 0.6204 -1.500 -0.2244 0.05247 0.04569 0.0185 0.9999 0.6723 -1.250 -0.2244 0.04938 0.04270 0.0240 0.9999 0.7157 -1.000 -0.2201 0.04641 0.03984 0.0282 0.9999 0.7500 -0.750 -0.2110 0.04388 0.03736 0.0310 0.9999 0.7842 -0.500 -0.1967 0.04156 0.03503 0.0326 0.9999 0.8205 -0.250 -0.1704 0.03980 0.03329 0.0310 0.9999 0.8432 0.000 -0.1312 0.03865 0.03210 0.0261 0.9999 0.8621 0.250 -0.0724 0.03802 0.03146 0.0164 0.9999 0.8709 0.500 -0.0026 0.03939 0.03268 0.0028 0.9999 0.8453 0.750 0.0826 0.04334 0.03637 -0.0170 0.9999 0.7324 1.000 0.1602 0.04675 0.03941 -0.0349 0.9999 0.5822 1.250 0.2206 0.04870 0.04095 -0.0463 0.9999 0.4546 1.500 0.2708 0.04945 0.04129 -0.0535 0.9999 0.3677 1.750 0.3154 0.04973 0.04137 -0.0583 0.9999 0.3199 2.000 0.3529 0.05025 0.04179 -0.0619 0.9999 0.2961 2.250 0.7860 0.03757 0.02458 -0.1048 0.2025 0.2691 2.500 0.8258 0.03876 0.02551 -0.1069 0.1930 0.2805 2.750 0.8855 0.03976 0.02645 -0.1121 0.1865 0.3045 3.000 1.0296 0.04066 0.02798 -0.1319 0.1724 1.0001 3.250 1.1061 0.04264 0.02926 -0.1384 0.1698 1.0001 3.500 1.1552 0.04440 0.03106 -0.1408 0.1706 1.0001 3.750 1.1967 0.04657 0.03343 -0.1418 0.1719 1.0001 4.000 1.2276 0.04848 0.03573 -0.1411 0.1754 1.0001 4.250 1.2569 0.05114 0.03882 -0.1403 0.1795 1.0001 4.500 1.2833 0.05405 0.04207 -0.1394 0.1838 1.0001 4.750 1.3055 0.05736 0.04569 -0.1378 0.1854 1.0001 5.000 1.3313 0.06120 0.04957 -0.1376 0.1896 1.0001