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 = 0.026 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.2436 0.04662 0.03692 -0.0235 1.0000 0.3255 -2.750 -0.2225 0.04559 0.03559 -0.0244 1.0000 0.3447 -2.500 -0.2064 0.04390 0.03395 -0.0239 1.0000 0.3675 -2.250 -0.1831 0.04283 0.03265 -0.0248 1.0000 0.3822 -2.000 -0.1553 0.04188 0.03126 -0.0268 1.0000 0.3928 -1.750 -0.1282 0.04085 0.02970 -0.0283 1.0000 0.4070 -1.500 -0.1051 0.04020 0.02901 -0.0288 1.0000 0.4284 -1.250 -0.0797 0.03980 0.02842 -0.0300 1.0000 0.4555 -1.000 -0.0584 0.03953 0.02831 -0.0302 1.0000 0.5009 -0.750 -0.0403 0.03896 0.02837 -0.0292 1.0000 0.5695 -0.500 -0.0245 0.03803 0.02853 -0.0268 1.0000 0.6945 -0.250 -0.0068 0.03770 0.02797 -0.0281 1.0000 1.0000 0.000 0.0146 0.03904 0.02849 -0.0295 1.0000 1.0000 0.250 0.0313 0.04041 0.02942 -0.0300 1.0000 1.0000 0.500 0.0467 0.04187 0.03056 -0.0306 1.0000 1.0000 0.750 0.0977 0.04403 0.03230 -0.0377 0.9813 1.0000 1.000 0.1739 0.04658 0.03442 -0.0485 0.9439 1.0000 1.250 0.2264 0.04859 0.03620 -0.0548 0.9113 1.0000 1.500 0.2670 0.05092 0.03836 -0.0592 0.8878 1.0000 1.750 0.3206 0.05292 0.04021 -0.0648 0.8570 1.0000 2.000 0.3441 0.05460 0.04180 -0.0655 0.8263 1.0000 2.250 0.4017 0.05621 0.04328 -0.0698 0.7888 1.0000 2.500 0.4280 0.05792 0.04493 -0.0707 0.7631 1.0000 2.750 0.4633 0.06025 0.04720 -0.0726 0.7422 1.0000 3.000 0.4593 0.06272 0.04965 -0.0709 0.7256 1.0000 3.250 0.5173 0.06463 0.05153 -0.0745 0.7039 1.0000 3.500 0.5109 0.06776 0.05461 -0.0728 0.6907 1.0000 3.750 0.5099 0.07011 0.05692 -0.0712 0.6721 1.0000 4.000 0.5597 0.07122 0.05792 -0.0718 0.6406 1.0000 4.250 0.5594 0.07344 0.06010 -0.0700 0.6201 1.0000 4.500 0.6404 0.07427 0.06085 -0.0724 0.5986 1.0000 4.750 0.6090 0.07888 0.06545 -0.0708 0.5937 1.0000 5.000 0.5947 0.08288 0.06944 -0.0702 0.5888 1.0000