XFOIL Version 6.94 Calculated polar for: Eiffel 385 (S.T. Ae) 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.080 Re = 0.450 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- 0.000 0.7988 0.01045 0.00355 -0.1075 0.5125 0.2350 0.500 0.8504 0.01059 0.00366 -0.1069 0.4931 0.2397 1.000 0.8995 0.01055 0.00365 -0.1056 0.4838 0.2528 1.500 0.9455 0.01071 0.00368 -0.1039 0.4569 0.2628 2.000 0.9789 0.01110 0.00392 -0.0998 0.4018 0.2908 2.500 1.0221 0.01119 0.00424 -0.0978 0.3810 0.3944 3.000 1.0538 0.01171 0.00454 -0.0935 0.3492 0.4110 3.500 1.0686 0.01220 0.00495 -0.0859 0.3117 0.4534 4.000 1.1066 0.01250 0.00543 -0.0830 0.2935 0.4998 4.500 1.1695 0.01223 0.00615 -0.0855 0.2876 1.0000 5.000 1.1903 0.01315 0.00679 -0.0798 0.2465 1.0000 5.500 1.1572 0.01577 0.00869 -0.0658 0.1511 1.0000 6.000 1.1430 0.01908 0.01155 -0.0575 0.0802 1.0000 6.500 1.1675 0.02076 0.01323 -0.0544 0.0688 1.0000 7.000 1.1832 0.02317 0.01575 -0.0509 0.0503 1.0000 7.500 1.2125 0.02453 0.01716 -0.0490 0.0468 1.0000 8.500 1.2540 0.02870 0.02154 -0.0448 0.0439 1.0000 9.000 1.2758 0.03079 0.02372 -0.0430 0.0412 1.0000 9.500 1.2876 0.03379 0.02681 -0.0411 0.0369 1.0000 10.000 1.2894 0.03798 0.03106 -0.0395 0.0308 1.0000 10.500 1.2917 0.04242 0.03554 -0.0383 0.0242 1.0000 11.000 1.2790 0.04895 0.04224 -0.0378 0.0223 1.0000 12.500 1.3089 0.06066 0.05446 -0.0360 0.0226 1.0000 13.000 1.2802 0.07016 0.06418 -0.0375 0.0219 1.0000 13.500 1.2871 0.07429 0.06828 -0.0373 0.0198 1.0000 14.000 1.2920 0.07882 0.07271 -0.0374 0.0174 1.0000 14.500 1.2835 0.08566 0.07958 -0.0383 0.0151 1.0000 15.000 1.2382 0.09878 0.09297 -0.0421 0.0147 1.0000 15.500 1.2339 0.10490 0.09922 -0.0429 0.0120 1.0000