XFOIL Version 6.96 Calculated polar for: FX 61-147 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.050 e 6 Ncrit = 5.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -10.000 -0.3587 0.10265 0.09587 -0.0491 1.0000 0.0459 -9.750 -0.3659 0.09781 0.09113 -0.0507 1.0000 0.0448 -9.500 -0.3771 0.09210 0.08557 -0.0532 1.0000 0.0435 -9.000 -0.4489 0.07951 0.07299 -0.0597 1.0000 0.0401 -8.750 -0.4674 0.07690 0.07045 -0.0580 1.0000 0.0399 -8.500 -0.4868 0.07429 0.06789 -0.0563 1.0000 0.0397 -8.250 -0.5041 0.07148 0.06508 -0.0548 1.0000 0.0396 -8.000 -0.5192 0.06864 0.06219 -0.0532 1.0000 0.0394 -7.750 -0.5310 0.06579 0.05926 -0.0516 1.0000 0.0393 -7.500 -0.5404 0.06279 0.05611 -0.0502 1.0000 0.0392 -7.250 -0.5455 0.05984 0.05299 -0.0489 1.0000 0.0391 -7.000 -0.5449 0.05654 0.04941 -0.0485 0.9993 0.0391 -6.750 -0.5174 0.05166 0.04393 -0.0531 0.9909 0.0393 -6.500 -0.4871 0.04749 0.03911 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0.2450 1.0000 13.000 1.2029 0.07531 0.06950 -0.0534 0.2254 1.0000 13.250 1.1946 0.07956 0.07382 -0.0543 0.2073 1.0000 13.500 1.1938 0.08256 0.07679 -0.0548 0.1873 1.0000 13.750 1.1904 0.08616 0.08032 -0.0556 0.1689 1.0000 14.000 1.1851 0.09026 0.08437 -0.0566 0.1519 1.0000 14.250 1.1804 0.09437 0.08838 -0.0577 0.1370 1.0000 14.500 1.1753 0.09869 0.09263 -0.0590 0.1239 1.0000 14.750 1.1705 0.10310 0.09698 -0.0604 0.1123 1.0000