XFOIL Version 6.94 Calculated polar for: ARA-D BL 12% 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 ------- -------- --------- --------- -------- ------- ------- -3.000 -0.0406 0.03888 0.02979 -0.0312 0.9997 0.3284 -2.750 -0.0163 0.03789 0.02870 -0.0311 0.9997 0.3431 -2.500 0.0093 0.03726 0.02793 -0.0311 0.9997 0.3620 -2.250 0.0334 0.03664 0.02742 -0.0308 0.9997 0.3848 -2.000 0.0586 0.03618 0.02717 -0.0307 0.9997 0.4135 -1.750 0.0834 0.03576 0.02728 -0.0304 0.9997 0.4504 -1.500 0.2611 0.02815 0.01914 -0.0371 0.4935 1.0003 -1.250 0.2891 0.02965 0.01940 -0.0358 0.4408 1.0003 -1.000 0.3187 0.03101 0.01994 -0.0352 0.4116 1.0003 -0.750 0.3490 0.03230 0.02066 -0.0350 0.3898 1.0003 -0.500 0.3799 0.03355 0.02157 -0.0350 0.3723 1.0003 -0.250 0.4117 0.03492 0.02273 -0.0353 0.3602 1.0003 0.000 0.4435 0.03636 0.02398 -0.0355 0.3517 1.0003 0.250 0.4747 0.03803 0.02547 -0.0358 0.3456 1.0003 0.500 0.5065 0.03976 0.02733 -0.0366 0.3404 1.0003 0.750 0.5370 0.04161 0.02919 -0.0372 0.3346 1.0003 1.000 0.5661 0.04354 0.03095 -0.0374 0.3287 1.0003 1.250 0.5949 0.04585 0.03328 -0.0381 0.3244 1.0003 1.500 0.6235 0.04842 0.03606 -0.0393 0.3228 1.0003 1.750 0.6508 0.05133 0.03916 -0.0405 0.3229 1.0003 2.000 0.6766 0.05458 0.04258 -0.0419 0.3239 1.0003 2.250 0.7004 0.05810 0.04625 -0.0433 0.3250 1.0003 2.500 0.7221 0.06189 0.05017 -0.0447 0.3259 1.0003 2.750 0.7236 0.06916 0.05820 -0.0503 0.3353 1.0003 3.000 0.7288 0.07519 0.06439 -0.0533 0.3422 1.0003 3.250 0.7441 0.07993 0.06914 -0.0547 0.3482 1.0003 3.500 0.6308 0.10060 0.09045 -0.0733 0.4328 1.0003 4.500 0.4837 0.12481 0.11488 -0.0919 0.7079 1.0003 4.750 0.5227 0.13198 0.12192 -0.0958 0.6997 1.0003 5.000 0.5144 0.13199 0.12184 -0.0929 0.6802 1.0003