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 = 0.026 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.1932 0.04502 0.03664 -0.0152 1.0000 0.3328 -2.750 -0.1713 0.04290 0.03445 -0.0154 1.0000 0.3413 -2.500 -0.1480 0.04125 0.03268 -0.0156 1.0000 0.3533 -2.250 -0.1226 0.03969 0.03096 -0.0162 1.0000 0.3674 -2.000 -0.0959 0.03854 0.02960 -0.0168 1.0000 0.3861 -1.750 -0.0715 0.03743 0.02859 -0.0167 1.0000 0.4081 -1.500 -0.0454 0.03656 0.02780 -0.0169 1.0000 0.4359 -1.250 -0.0203 0.03582 0.02740 -0.0169 1.0000 0.4701 -0.750 0.2636 0.02967 0.02015 -0.0325 0.4861 1.0000 -0.500 0.2912 0.03096 0.02069 -0.0315 0.4550 1.0000 -0.250 0.3194 0.03224 0.02144 -0.0309 0.4331 1.0000 0.000 0.3475 0.03353 0.02226 -0.0304 0.4151 1.0000 0.250 0.3761 0.03493 0.02348 -0.0303 0.3996 1.0000 0.500 0.4043 0.03643 0.02471 -0.0301 0.3883 1.0000 0.750 0.4330 0.03817 0.02647 -0.0306 0.3806 1.0000 1.000 0.4614 0.04001 0.02821 -0.0308 0.3748 1.0000 1.250 0.4894 0.04204 0.03009 -0.0310 0.3703 1.0000 1.500 0.5151 0.04456 0.03277 -0.0320 0.3670 1.0000 1.750 0.5383 0.04734 0.03573 -0.0331 0.3633 1.0000 2.000 0.5594 0.05034 0.03885 -0.0341 0.3600 1.0000 2.250 0.5779 0.05373 0.04236 -0.0353 0.3586 1.0000 2.500 0.5914 0.05799 0.04679 -0.0372 0.3610 1.0000 2.750 0.6024 0.06252 0.05142 -0.0391 0.3651 1.0000 3.000 0.6165 0.06671 0.05559 -0.0403 0.3689 1.0000 3.250 0.5857 0.07580 0.06506 -0.0471 0.3871 1.0000 3.750 0.5078 0.09252 0.08196 -0.0599 0.4499 1.0000 4.250 0.3292 0.10421 0.09417 -0.0746 0.8110 1.0000 4.500 0.3464 0.10794 0.09775 -0.0761 0.8057 1.0000 4.750 0.3716 0.11328 0.10291 -0.0788 0.7995 1.0000 5.000 0.3759 0.11458 0.10411 -0.0774 0.7805 1.0000