XFOIL Version 6.94 Calculated polar for: nabab 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.005 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.2115 0.05787 0.03865 -0.0173 0.9999 1.0001 -2.750 -0.2123 0.05546 0.03553 -0.0194 0.9999 1.0001 -2.500 -0.1855 0.05442 0.03216 -0.0228 0.9999 1.0001 -2.250 -0.1565 0.05414 0.02993 -0.0231 0.9999 1.0001 -2.000 -0.1312 0.05397 0.02838 -0.0223 0.9999 1.0001 -1.750 -0.1073 0.05384 0.02717 -0.0212 0.9999 1.0001 -1.500 -0.0842 0.05374 0.02620 -0.0201 0.9999 1.0001 -1.250 -0.0614 0.05368 0.02539 -0.0190 0.9999 1.0001 -1.000 -0.0388 0.05366 0.02473 -0.0179 0.9999 1.0001 -0.750 -0.0163 0.05366 0.02419 -0.0168 0.9999 1.0001 -0.500 0.0061 0.05369 0.02376 -0.0157 0.9999 1.0001 -0.250 0.0285 0.05376 0.02345 -0.0146 0.9999 1.0001 0.000 0.0508 0.05385 0.02325 -0.0136 0.9999 1.0001 0.250 0.0732 0.05398 0.02315 -0.0126 0.9999 1.0001 0.500 0.0955 0.05414 0.02316 -0.0116 0.9999 1.0001 0.750 0.1177 0.05433 0.02327 -0.0107 0.9999 1.0001 1.000 0.1399 0.05455 0.02350 -0.0098 0.9999 1.0001 1.250 0.1621 0.05482 0.02384 -0.0089 0.9999 1.0001 1.500 0.1841 0.05513 0.02431 -0.0081 0.9999 1.0001 1.750 0.2060 0.05550 0.02489 -0.0074 0.9999 1.0001 2.000 0.2275 0.05593 0.02562 -0.0068 0.9999 1.0001 2.250 0.2485 0.05644 0.02651 -0.0062 0.9999 1.0001 2.500 0.2688 0.05708 0.02760 -0.0058 0.9999 1.0001 2.750 0.2877 0.05789 0.02892 -0.0055 0.9999 1.0001 3.000 0.3042 0.05898 0.03055 -0.0054 0.9999 1.0001 3.250 0.3164 0.06052 0.03259 -0.0056 0.9999 1.0001 3.500 0.3220 0.06269 0.03506 -0.0059 0.9999 1.0001 3.750 0.3216 0.06548 0.03787 -0.0064 0.9999 1.0001 4.000 0.3204 0.06843 0.04068 -0.0069 0.9999 1.0001 4.250 0.3214 0.07126 0.04334 -0.0076 0.9999 1.0001 4.500 0.3249 0.07394 0.04579 -0.0082 0.9999 1.0001 4.750 0.3299 0.07652 0.04826 -0.0088 0.9999 1.0001 5.000 0.3362 0.07903 0.05069 -0.0095 0.9999 1.0001