XFOIL Version 6.94 Calculated polar for: nabac 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.008 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.3987 0.05875 0.04544 0.0230 0.9999 0.5144 -2.750 -0.3680 0.05571 0.04233 0.0210 0.9999 0.5314 -2.500 -0.3359 0.05265 0.03952 0.0198 0.9999 0.5579 -2.250 -0.3027 0.04942 0.03689 0.0195 0.9999 0.5999 -2.000 -0.2601 0.04545 0.03429 0.0189 0.9999 0.6895 -1.750 -0.0900 0.04175 0.02856 -0.0106 0.9999 1.0001 -1.500 -0.0566 0.04171 0.02752 -0.0098 0.9999 1.0001 -1.250 -0.0262 0.04164 0.02698 -0.0087 0.9999 1.0001 -1.000 0.0043 0.04157 0.02672 -0.0078 0.9999 1.0001 -0.750 0.0357 0.04151 0.02668 -0.0074 0.9999 1.0001 -0.500 0.0686 0.04154 0.02691 -0.0079 0.9999 1.0001 -0.250 0.1016 0.04185 0.02767 -0.0099 0.9999 1.0001 0.000 0.1197 0.04325 0.02963 -0.0128 0.9999 1.0001 0.250 0.1080 0.04610 0.03236 -0.0126 0.9999 1.0001 0.500 0.2579 0.05054 0.03519 -0.0410 0.8208 1.0001 0.750 0.3075 0.05291 0.03675 -0.0458 0.7697 1.0001 1.000 0.3394 0.05532 0.03857 -0.0481 0.7409 1.0001 1.250 0.3686 0.05780 0.04051 -0.0500 0.7209 1.0001 1.500 0.3881 0.06048 0.04282 -0.0515 0.7081 1.0001 1.750 0.4126 0.06320 0.04510 -0.0531 0.6976 1.0001 2.000 0.4240 0.06615 0.04776 -0.0543 0.6926 1.0001 2.250 0.4343 0.06915 0.05047 -0.0552 0.6895 1.0001 2.500 0.4427 0.07221 0.05323 -0.0559 0.6883 1.0001 2.750 0.4492 0.07529 0.05597 -0.0564 0.6886 1.0001 3.000 0.4537 0.07839 0.05879 -0.0568 0.6906 1.0001 3.250 0.4585 0.08151 0.06164 -0.0571 0.6936 1.0001 3.500 0.4650 0.08467 0.06453 -0.0575 0.6973 1.0001 3.750 0.4548 0.08744 0.06708 -0.0560 0.7044 1.0001 4.000 0.4497 0.09028 0.06967 -0.0551 0.7121 1.0001 4.250 0.4545 0.09340 0.07254 -0.0555 0.7195 1.0001 4.500 0.4526 0.09620 0.07514 -0.0551 0.7288 1.0001 4.750 0.4478 0.09886 0.07758 -0.0544 0.7404 1.0001 5.000 0.4561 0.10215 0.08068 -0.0554 0.7510 1.0001