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 = 0.030 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.2368 0.04549 0.03598 -0.0263 1.0000 0.2930 -2.750 -0.2127 0.04426 0.03431 -0.0277 1.0000 0.2991 -2.500 -0.1963 0.04216 0.03238 -0.0275 1.0000 0.3202 -2.250 -0.1761 0.04177 0.03183 -0.0277 1.0000 0.3436 -2.000 -0.1508 0.04071 0.03042 -0.0291 1.0000 0.3537 -1.750 -0.1271 0.03971 0.02908 -0.0298 1.0000 0.3669 -1.500 -0.1032 0.03935 0.02850 -0.0306 1.0000 0.3845 -1.250 -0.0782 0.03904 0.02796 -0.0317 1.0000 0.4057 -1.000 -0.0562 0.03883 0.02784 -0.0322 1.0000 0.4355 -0.750 -0.0353 0.03872 0.02799 -0.0324 1.0000 0.4857 -0.500 -0.0146 0.03840 0.02834 -0.0323 1.0000 0.5664 -0.250 -0.0106 0.03699 0.02865 -0.0274 1.0000 0.9039 0.000 0.0573 0.03854 0.02854 -0.0378 0.9802 1.0000 0.250 0.1183 0.04035 0.02971 -0.0457 0.9561 1.0000 0.500 0.1633 0.04177 0.03079 -0.0504 0.9288 1.0000 0.750 0.2480 0.04380 0.03241 -0.0609 0.8957 1.0000 1.000 0.2701 0.04488 0.03336 -0.0615 0.8696 1.0000 1.250 0.3347 0.04661 0.03489 -0.0682 0.8447 1.0000 1.500 0.3549 0.04774 0.03594 -0.0680 0.8162 1.0000 1.750 0.4030 0.04884 0.03691 -0.0709 0.7849 1.0000 2.000 0.4870 0.04857 0.03653 -0.0765 0.7523 1.0000 2.250 0.4706 0.05120 0.03912 -0.0726 0.7295 1.0000 2.500 0.5288 0.05163 0.03950 -0.0754 0.7062 1.0000 2.750 0.5316 0.05439 0.04224 -0.0739 0.6871 1.0000 3.000 0.5399 0.05661 0.04443 -0.0728 0.6672 1.0000 3.250 0.6008 0.05665 0.04434 -0.0738 0.6450 1.0000 3.500 0.5817 0.05998 0.04766 -0.0712 0.6229 1.0000 3.750 0.6646 0.05730 0.04479 -0.0705 0.5972 1.0000 4.000 0.6300 0.06251 0.05006 -0.0689 0.5787 1.0000 4.250 0.6496 0.06442 0.05193 -0.0683 0.5637 1.0000 4.500 0.7242 0.06281 0.05017 -0.0673 0.5493 1.0000 4.750 0.6746 0.07026 0.05771 -0.0673 0.5408 1.0000 5.000 0.6469 0.07561 0.06305 -0.0665 0.5338 1.0000