XFOIL Version 6.94 Calculated polar for: ARA-D BL 13% 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 -0.1928 0.04137 0.03411 -0.0190 1.0000 0.2213 -2.750 -0.1697 0.04003 0.03267 -0.0192 1.0000 0.2281 -2.500 -0.0258 0.03498 0.02720 -0.0377 0.9497 0.2558 -2.250 0.1066 0.02888 0.02114 -0.0459 0.8099 0.2971 -2.000 0.1170 0.02874 0.01912 -0.0356 0.4563 0.3117 -1.750 0.1387 0.02939 0.01866 -0.0336 0.3678 0.3318 -1.500 0.1668 0.02947 0.01836 -0.0327 0.3371 0.3585 -1.250 0.1957 0.02939 0.01820 -0.0319 0.3177 0.3926 -1.000 0.2254 0.02926 0.01816 -0.0312 0.3044 0.4443 -0.750 0.2509 0.02833 0.01826 -0.0293 0.2958 0.5630 -0.500 0.3010 0.02804 0.01830 -0.0298 0.2865 1.0000 -0.250 0.3314 0.02910 0.01899 -0.0295 0.2815 1.0000 0.000 0.3608 0.03011 0.01984 -0.0292 0.2768 1.0000 0.250 0.3894 0.03123 0.02079 -0.0288 0.2716 1.0000 0.500 0.4173 0.03247 0.02182 -0.0285 0.2661 1.0000 0.750 0.4445 0.03416 0.02323 -0.0283 0.2614 1.0000 1.000 0.4720 0.03578 0.02482 -0.0281 0.2590 1.0000 1.250 0.4995 0.03733 0.02645 -0.0280 0.2578 1.0000 1.500 0.5263 0.03909 0.02831 -0.0279 0.2570 1.0000 1.750 0.5523 0.04107 0.03039 -0.0278 0.2568 1.0000 2.000 0.5770 0.04322 0.03267 -0.0278 0.2565 1.0000 2.250 0.6003 0.04552 0.03514 -0.0278 0.2559 1.0000 2.500 0.6220 0.04803 0.03782 -0.0278 0.2552 1.0000 2.750 0.6417 0.05087 0.04083 -0.0279 0.2553 1.0000 3.000 0.6605 0.05416 0.04424 -0.0281 0.2571 1.0000 3.250 0.6807 0.05778 0.04782 -0.0282 0.2594 1.0000 3.500 0.6657 0.06558 0.05667 -0.0319 0.2798 1.0000 3.750 0.6938 0.06910 0.05995 -0.0316 0.2845 1.0000 4.000 0.4385 0.10703 0.09950 -0.0734 0.5552 1.0000 4.500 0.4372 0.11085 0.10310 -0.0704 0.5191 1.0000 5.000 0.4751 0.11711 0.10910 -0.0695 0.4870 1.0000