XFOIL Version 6.94 Calculated polar for: Fage & Collins 3 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.080 Re = 0.450 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- 0.000 0.4312 0.01040 0.00301 -0.0719 0.6240 0.0477 0.500 0.4557 0.01064 0.00275 -0.0649 0.5257 0.0475 1.000 0.4893 0.01085 0.00262 -0.0601 0.4662 0.0482 1.500 0.5285 0.01105 0.00259 -0.0566 0.4292 0.0486 2.000 0.5705 0.01122 0.00260 -0.0537 0.4028 0.0544 2.500 0.8054 0.01032 0.00365 -0.0961 0.3658 1.0000 3.000 0.8496 0.01068 0.00389 -0.0938 0.3521 1.0000 3.500 0.8937 0.01105 0.00420 -0.0916 0.3404 1.0000 4.000 0.9377 0.01146 0.00454 -0.0894 0.3308 1.0000 4.500 0.9815 0.01191 0.00494 -0.0872 0.3217 1.0000 5.000 1.0262 0.01228 0.00536 -0.0852 0.3129 1.0000 5.500 1.0698 0.01269 0.00581 -0.0830 0.3035 1.0000 6.000 1.1121 0.01305 0.00619 -0.0806 0.2906 1.0000 6.500 1.1534 0.01341 0.00659 -0.0780 0.2764 1.0000 7.000 1.1940 0.01377 0.00702 -0.0753 0.2598 1.0000 7.500 1.2311 0.01425 0.00742 -0.0720 0.2201 1.0000 8.000 1.2166 0.01784 0.00984 -0.0600 0.0313 1.0000 8.500 1.2374 0.01927 0.01138 -0.0539 0.0200 1.0000 9.000 1.2539 0.02098 0.01331 -0.0475 0.0167 1.0000 9.500 1.2608 0.02325 0.01585 -0.0400 0.0151 1.0000 10.000 1.2669 0.02567 0.01850 -0.0333 0.0138 1.0000 10.500 1.2695 0.02857 0.02159 -0.0273 0.0128 1.0000 11.000 1.2650 0.03234 0.02556 -0.0218 0.0123 1.0000 11.500 1.2485 0.03769 0.03110 -0.0168 0.0116 1.0000 12.000 1.2342 0.04369 0.03734 -0.0134 0.0108 1.0000 12.500 1.2315 0.04896 0.04275 -0.0110 0.0109 1.0000 13.000 1.2406 0.05321 0.04709 -0.0080 0.0111 1.0000 13.500 1.2423 0.05860 0.05277 -0.0078 0.0103 1.0000 14.000 1.2601 0.06250 0.05687 -0.0043 0.0106 1.0000 14.500 1.2583 0.06868 0.06335 -0.0040 0.0100 1.0000