XFOIL Version 6.94 Calculated polar for: ARA-D 13% AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.040 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.2332 0.04252 0.03374 -0.0288 1.0000 0.2384 -2.750 -0.2094 0.04125 0.03210 -0.0298 1.0000 0.2442 -2.500 -0.1866 0.04080 0.03128 -0.0304 1.0000 0.2501 -2.250 -0.1640 0.03838 0.02882 -0.0316 1.0000 0.2628 -2.000 -0.1470 0.03835 0.02885 -0.0315 1.0000 0.2843 -1.750 -0.1231 0.03859 0.02862 -0.0324 0.9987 0.3000 -1.500 -0.0658 0.03680 0.02669 -0.0388 0.9848 0.3227 -1.250 -0.0008 0.03619 0.02578 -0.0460 0.9696 0.3585 -1.000 0.0641 0.03535 0.02519 -0.0529 0.9520 0.4232 -0.750 0.0998 0.03445 0.02504 -0.0546 0.9329 0.5244 -0.500 0.1570 0.03309 0.02444 -0.0577 0.9118 1.0000 -0.250 0.2144 0.03396 0.02474 -0.0632 0.8914 1.0000 0.000 0.2582 0.03436 0.02484 -0.0661 0.8660 1.0000 0.250 0.3312 0.03458 0.02475 -0.0730 0.8454 1.0000 0.500 0.3500 0.03539 0.02545 -0.0721 0.8212 1.0000 0.750 0.4219 0.03499 0.02483 -0.0771 0.8019 1.0000 1.000 0.4324 0.03600 0.02576 -0.0748 0.7742 1.0000 1.250 0.4911 0.03508 0.02467 -0.0762 0.7531 1.0000 1.500 0.5043 0.03617 0.02570 -0.0741 0.7254 1.0000 1.750 0.5463 0.03576 0.02516 -0.0737 0.7062 1.0000 2.000 0.5628 0.03712 0.02649 -0.0723 0.6821 1.0000 2.250 0.5948 0.03715 0.02643 -0.0711 0.6611 1.0000 2.500 0.6208 0.03782 0.02697 -0.0695 0.6407 1.0000 2.750 0.6439 0.03826 0.02734 -0.0678 0.6156 1.0000 3.000 0.6796 0.03759 0.02640 -0.0652 0.5977 1.0000 3.250 0.6943 0.03875 0.02760 -0.0639 0.5700 1.0000 3.500 0.7296 0.03791 0.02648 -0.0615 0.5534 1.0000 3.750 0.7416 0.04035 0.02902 -0.0611 0.5325 1.0000 4.000 0.7656 0.04111 0.02972 -0.0599 0.5150 1.0000 4.250 0.7994 0.04101 0.02936 -0.0582 0.5022 1.0000 4.500 0.8038 0.04456 0.03313 -0.0582 0.4857 1.0000 4.750 0.8189 0.04662 0.03525 -0.0575 0.4707 1.0000 5.000 0.8488 0.04689 0.03540 -0.0562 0.4588 1.0000