XFOIL Version 6.94 Calculated polar for: nabad 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.026 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.3724 0.05067 0.04024 -0.0018 0.9999 0.1665 -2.750 -0.3410 0.04744 0.03641 -0.0023 0.9999 0.1543 -2.500 -0.3111 0.04430 0.03307 -0.0023 0.9999 0.1467 -2.250 -0.2787 0.04166 0.02995 -0.0022 0.9999 0.1396 -2.000 -0.2429 0.03920 0.02705 -0.0025 0.9999 0.1338 -1.750 -0.2042 0.03706 0.02449 -0.0030 0.9999 0.1295 -1.500 -0.1647 0.03506 0.02235 -0.0037 0.9999 0.1269 -1.250 -0.1255 0.03333 0.02057 -0.0043 0.9999 0.1255 -1.000 -0.0885 0.03182 0.01917 -0.0045 0.9999 0.1260 -0.750 -0.0533 0.03053 0.01810 -0.0046 0.9999 0.1292 -0.500 -0.0189 0.02937 0.01735 -0.0048 0.9999 0.1372 -0.250 0.0405 0.02858 0.01645 -0.0079 0.6785 0.1675 0.000 0.1074 0.02701 0.01589 -0.0132 0.4996 1.0001 0.250 0.1352 0.02774 0.01578 -0.0121 0.4799 1.0001 0.500 0.1636 0.02841 0.01593 -0.0113 0.4675 1.0001 0.750 0.1920 0.02914 0.01619 -0.0106 0.4597 1.0001 1.000 0.2217 0.02968 0.01653 -0.0103 0.4547 1.0001 1.250 0.2515 0.03024 0.01693 -0.0100 0.4513 1.0001 1.500 0.2815 0.03084 0.01737 -0.0098 0.4494 1.0001 1.750 0.3116 0.03144 0.01785 -0.0098 0.4487 1.0001 2.000 0.3421 0.03206 0.01839 -0.0097 0.4491 1.0001 2.250 0.3697 0.03279 0.01904 -0.0095 0.4394 1.0001 2.500 0.3898 0.03397 0.01965 -0.0077 0.4014 1.0001 2.750 0.4113 0.03462 0.02003 -0.0066 0.3637 1.0001 3.000 0.4338 0.03519 0.02036 -0.0056 0.3373 1.0001 3.250 0.4564 0.03564 0.02060 -0.0047 0.3140 1.0001 3.500 0.4789 0.03608 0.02074 -0.0037 0.2938 1.0001 3.750 0.5035 0.03609 0.02086 -0.0031 0.2768 1.0001 4.000 0.5273 0.03634 0.02099 -0.0023 0.2612 1.0001 4.250 0.5518 0.03652 0.02116 -0.0015 0.2459 1.0001 4.500 0.5772 0.03685 0.02159 -0.0007 0.2317 1.0001 4.750 0.6023 0.03742 0.02216 0.0002 0.2157 1.0001 5.000 0.6279 0.03826 0.02316 0.0011 0.1958 1.0001