XFOIL Version 6.94 Calculated polar for: manu01 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.005 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.0613 0.07945 0.05803 -0.0195 0.9999 1.0001 -2.750 -0.0627 0.07748 0.05632 -0.0183 0.9999 1.0001 -2.500 -0.0652 0.07549 0.05459 -0.0170 0.9999 1.0001 -2.250 -0.0693 0.07345 0.05283 -0.0155 0.9999 1.0001 -2.000 -0.0748 0.07138 0.05104 -0.0138 0.9999 1.0001 -1.750 -0.0805 0.06932 0.04921 -0.0121 0.9999 1.0001 -1.500 -0.0832 0.06737 0.04734 -0.0112 0.9999 1.0001 -1.250 -0.0785 0.06577 0.04560 -0.0115 0.9999 1.0001 -1.000 -0.0624 0.06469 0.04411 -0.0138 0.9999 1.0001 -0.750 -0.0334 0.06420 0.04291 -0.0182 0.9999 1.0001 -0.500 0.0057 0.06428 0.04212 -0.0240 0.9999 1.0001 -0.250 0.0497 0.06479 0.04165 -0.0300 0.9999 1.0001 0.000 0.0924 0.06553 0.04143 -0.0350 0.9999 1.0001 0.250 0.1307 0.06635 0.04138 -0.0386 0.9999 1.0001 0.500 0.1641 0.06719 0.04161 -0.0410 0.9999 1.0001 0.750 0.1934 0.06811 0.04210 -0.0426 0.9999 1.0001 1.000 0.2189 0.06917 0.04294 -0.0438 0.9999 1.0001 1.250 0.2396 0.07057 0.04428 -0.0448 0.9999 1.0001 1.500 0.2529 0.07256 0.04632 -0.0456 0.9999 1.0001 1.750 0.2549 0.07556 0.04939 -0.0464 0.9999 1.0001 2.000 0.2492 0.07928 0.05299 -0.0471 0.9999 1.0001 2.250 0.2455 0.08293 0.05636 -0.0479 0.9999 1.0001 2.500 0.2462 0.08625 0.05934 -0.0488 0.9999 1.0001 2.750 0.2499 0.08934 0.06208 -0.0497 0.9999 1.0001 3.000 0.2554 0.09229 0.06467 -0.0507 0.9999 1.0001 3.250 0.2624 0.09511 0.06713 -0.0515 0.9999 1.0001 3.500 0.2703 0.09786 0.06953 -0.0524 0.9999 1.0001 3.750 0.2788 0.10058 0.07190 -0.0532 0.9999 1.0001 4.000 0.2880 0.10324 0.07423 -0.0540 0.9999 1.0001 4.250 0.2974 0.10587 0.07654 -0.0548 0.9999 1.0001 4.500 0.3073 0.10848 0.07884 -0.0555 0.9999 1.0001 4.750 0.3175 0.11106 0.08112 -0.0562 0.9999 1.0001 5.000 0.3279 0.11364 0.08341 -0.0569 0.9999 1.0001