XFOIL Version 6.94 Calculated polar for: WORTMANN FX 66-17A-175 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.080 Re = 0.300 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- 0.000 0.4751 0.01508 0.00727 -0.0943 0.5120 0.5660 0.500 0.5312 0.01545 0.00752 -0.0949 0.5077 0.5703 1.000 0.5857 0.01548 0.00755 -0.0950 0.5033 0.5737 1.500 0.6402 0.01554 0.00766 -0.0952 0.4989 0.5775 2.000 0.6954 0.01571 0.00783 -0.0954 0.4951 0.5816 2.500 0.7513 0.01591 0.00801 -0.0959 0.4913 0.5862 3.000 0.8067 0.01655 0.00860 -0.0964 0.4870 0.5911 3.500 0.8590 0.01678 0.00898 -0.0962 0.4843 0.5965 4.000 0.9105 0.01712 0.00949 -0.0960 0.4809 0.6019 4.500 0.9622 0.01755 0.01007 -0.0958 0.4774 0.6088 5.000 1.0147 0.01798 0.01061 -0.0958 0.4742 0.6154 5.500 1.0674 0.01835 0.01115 -0.0958 0.4713 0.6231 6.000 1.1210 0.01888 0.01179 -0.0960 0.4685 0.6326 6.500 1.1713 0.01990 0.01302 -0.0959 0.4653 0.6443 7.000 1.2150 0.02055 0.01402 -0.0945 0.4624 0.6593 7.500 1.2635 0.01982 0.01356 -0.0932 0.4531 0.6815 8.000 1.3158 0.01840 0.01234 -0.0920 0.4397 0.7355 9.000 1.4021 0.01637 0.01077 -0.0871 0.3998 1.0000 9.500 1.3910 0.01681 0.01100 -0.0753 0.3301 1.0000 11.500 1.1690 0.04641 0.03970 -0.0461 0.1004 1.0000 12.000 1.1161 0.05696 0.04996 -0.0448 0.0426 1.0000 12.500 1.0865 0.06592 0.05879 -0.0444 0.0143 1.0000 13.000 1.0751 0.07332 0.06632 -0.0444 0.0094 1.0000 13.500 1.0722 0.07995 0.07313 -0.0448 0.0078 1.0000 14.000 1.0672 0.08704 0.08036 -0.0454 0.0072 1.0000 14.500 1.0670 0.09365 0.08711 -0.0461 0.0067 1.0000 15.000 1.0721 0.09951 0.09307 -0.0468 0.0064 1.0000 15.500 1.0813 0.10464 0.09827 -0.0472 0.0061 1.0000 16.000 1.0998 0.10804 0.10169 -0.0468 0.0059 1.0000 16.500 1.1275 0.10989 0.10358 -0.0457 0.0058 1.0000 17.000 1.1603 0.11111 0.10497 -0.0440 0.0059 1.0000 17.500 1.1897 0.11356 0.10784 -0.0422 0.0065 1.0000 18.000 1.2146 0.11708 0.11165 -0.0414 0.0071 1.0000