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 = 0.030 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -2.750 -0.3223 0.06261 0.05513 0.0290 0.9999 0.6423 -2.500 -0.3273 0.05968 0.05234 0.0351 0.9999 0.6937 -2.250 -0.3295 0.05622 0.04893 0.0407 0.9999 0.7354 -2.000 -0.3286 0.05250 0.04535 0.0459 0.9999 0.7715 -1.750 -0.3252 0.04950 0.04243 0.0492 0.9999 0.8031 -1.500 -0.3174 0.04640 0.03941 0.0518 0.9999 0.8331 -1.250 -0.3105 0.04374 0.03679 0.0540 0.9999 0.8606 -1.000 -0.2968 0.04130 0.03441 0.0544 0.9999 0.8819 -0.750 -0.2731 0.03890 0.03205 0.0526 0.9999 0.9044 -0.500 -0.2270 0.03688 0.03005 0.0454 0.9999 0.9255 -0.250 -0.1805 0.03623 0.02939 0.0373 0.9999 0.9289 0.000 -0.1298 0.03686 0.02988 0.0279 0.9999 0.9128 0.250 -0.0295 0.04139 0.03401 0.0041 0.9999 0.8127 0.500 0.0684 0.04553 0.03778 -0.0196 0.9999 0.6800 0.750 0.1438 0.04814 0.04005 -0.0360 0.9999 0.5533 1.000 0.2037 0.04961 0.04110 -0.0468 0.9999 0.4519 1.250 0.2511 0.05028 0.04147 -0.0533 0.9999 0.3886 1.500 0.2943 0.05084 0.04169 -0.0583 0.9999 0.3490 1.750 0.3356 0.05116 0.04188 -0.0623 0.9999 0.3283 2.000 0.3741 0.05149 0.04220 -0.0662 0.9999 0.3155 2.250 0.4173 0.05211 0.04271 -0.0703 0.9999 0.3010 2.500 0.8280 0.03944 0.02625 -0.1080 0.2252 0.3225 2.750 0.8808 0.04075 0.02747 -0.1124 0.2129 0.3634 3.000 0.9905 0.04168 0.02900 -0.1256 0.2022 1.0001 3.250 1.0411 0.04313 0.02967 -0.1273 0.1968 1.0001 3.500 1.1115 0.04535 0.03149 -0.1332 0.1909 1.0001 3.750 1.1692 0.04772 0.03399 -0.1368 0.1890 1.0001 4.000 1.2153 0.05035 0.03673 -0.1391 0.1903 1.0001 4.250 1.2459 0.05235 0.03919 -0.1382 0.1928 1.0001 4.500 1.2702 0.05468 0.04212 -0.1368 0.1970 1.0001 4.750 1.2935 0.05794 0.04590 -0.1353 0.2017 1.0001 5.000 1.3158 0.06109 0.04934 -0.1343 0.2062 1.0001