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 = 0.045 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.2308 0.04065 0.03208 -0.0300 1.0000 0.2180 -2.750 -0.2090 0.03944 0.03066 -0.0306 1.0000 0.2246 -2.500 -0.1869 0.03855 0.02953 -0.0312 1.0000 0.2331 -2.250 -0.1497 0.03774 0.02836 -0.0342 0.9944 0.2442 -2.000 -0.0908 0.03628 0.02672 -0.0412 0.9799 0.2773 -1.750 -0.0357 0.03564 0.02570 -0.0468 0.9652 0.2975 -1.500 0.0266 0.03420 0.02417 -0.0532 0.9520 0.3296 -1.250 0.0660 0.03360 0.02351 -0.0556 0.9340 0.3625 -1.000 0.1206 0.03266 0.02294 -0.0603 0.9153 0.4399 -0.750 0.1998 0.03012 0.02202 -0.0663 0.9013 1.0000 -0.500 0.2206 0.03072 0.02222 -0.0655 0.8784 1.0000 -0.250 0.2907 0.03055 0.02162 -0.0720 0.8583 1.0000 0.000 0.3286 0.03099 0.02183 -0.0734 0.8375 1.0000 0.250 0.3660 0.03123 0.02188 -0.0746 0.8161 1.0000 0.500 0.4240 0.03090 0.02132 -0.0773 0.7997 1.0000 0.750 0.4370 0.03162 0.02197 -0.0751 0.7726 1.0000 1.000 0.4822 0.03096 0.02111 -0.0749 0.7535 1.0000 1.250 0.5004 0.03170 0.02179 -0.0732 0.7274 1.0000 1.500 0.5339 0.03145 0.02140 -0.0719 0.7081 1.0000 1.750 0.5572 0.03220 0.02207 -0.0706 0.6864 1.0000 2.000 0.5834 0.03241 0.02222 -0.0691 0.6640 1.0000 2.250 0.6135 0.03246 0.02205 -0.0671 0.6462 1.0000 2.500 0.6335 0.03312 0.02269 -0.0657 0.6189 1.0000 2.750 0.6656 0.03250 0.02175 -0.0629 0.6011 1.0000 3.000 0.6839 0.03341 0.02270 -0.0618 0.5731 1.0000 3.250 0.7144 0.03286 0.02186 -0.0595 0.5548 1.0000 3.500 0.7342 0.03428 0.02330 -0.0589 0.5328 1.0000 3.750 0.7595 0.03472 0.02365 -0.0576 0.5141 1.0000 4.000 0.7891 0.03519 0.02377 -0.0561 0.4997 1.0000 4.250 0.8046 0.03718 0.02602 -0.0561 0.4810 1.0000 4.500 0.8284 0.03800 0.02679 -0.0551 0.4656 1.0000 4.750 0.8584 0.03838 0.02691 -0.0539 0.4534 1.0000 5.000 0.8721 0.04109 0.02982 -0.0537 0.4402 1.0000