XFOIL Version 6.94 Calculated polar for: manu4/36 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.008 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.2851 0.07309 0.05763 0.0294 0.9999 0.8352 -2.750 -0.2961 0.07062 0.05534 0.0325 0.9999 0.8319 -2.500 -0.2960 0.06806 0.05292 0.0335 0.9999 0.8320 -2.250 -0.2862 0.06554 0.05053 0.0329 0.9999 0.8354 -2.000 -0.2633 0.06311 0.04818 0.0300 0.9999 0.8441 -1.750 -0.2254 0.06084 0.04597 0.0242 0.9999 0.8619 -1.500 -0.0671 0.05880 0.04391 -0.0064 0.9999 0.9733 -1.250 -0.0623 0.05717 0.04243 -0.0076 0.9999 1.0001 -1.000 -0.0648 0.05528 0.04063 -0.0065 0.9999 1.0001 -0.750 -0.0549 0.05390 0.03907 -0.0076 0.9999 1.0001 -0.500 -0.0269 0.05323 0.03797 -0.0119 0.9999 1.0001 -0.250 0.0187 0.05331 0.03737 -0.0193 0.9999 1.0001 0.000 0.0771 0.05405 0.03720 -0.0286 0.9999 1.0001 0.250 0.1328 0.05512 0.03733 -0.0364 0.9999 1.0001 0.500 0.1745 0.05624 0.03791 -0.0409 0.9999 1.0001 0.750 0.2024 0.05772 0.03921 -0.0436 0.9999 1.0001 1.000 0.2109 0.06038 0.04196 -0.0451 0.9999 1.0001 1.250 0.1970 0.06473 0.04634 -0.0458 0.9999 1.0001 1.500 0.1897 0.06870 0.05006 -0.0468 0.9999 1.0001 1.750 0.1905 0.07213 0.05312 -0.0480 0.9999 1.0001 2.000 0.1956 0.07521 0.05578 -0.0493 0.9999 1.0001 2.250 0.2028 0.07811 0.05827 -0.0504 0.9999 1.0001 2.500 0.2112 0.08090 0.06067 -0.0515 0.9999 1.0001 2.750 0.2208 0.08359 0.06296 -0.0524 0.9999 1.0001 3.000 0.2308 0.08624 0.06524 -0.0534 0.9999 1.0001 3.250 0.2413 0.08885 0.06749 -0.0542 0.9999 1.0001 3.500 0.2521 0.09144 0.06974 -0.0550 0.9999 1.0001 3.750 0.2631 0.09401 0.07199 -0.0558 0.9999 1.0001 4.000 0.2743 0.09658 0.07424 -0.0565 0.9999 1.0001 4.250 0.2856 0.09915 0.07651 -0.0572 0.9999 1.0001 4.500 0.2970 0.10171 0.07880 -0.0579 0.9999 1.0001 4.750 0.3085 0.10427 0.08109 -0.0586 0.9999 1.0001 5.000 0.3200 0.10684 0.08342 -0.0592 0.9999 1.0001