XFOIL Version 6.96 Calculated polar for: FX 66-S-196 V1 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 1.000 e 6 Ncrit = 5.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -18.250 -0.7780 0.09188 0.08913 -0.0680 0.9484 0.0059 -18.000 -0.7816 0.08248 0.07950 -0.0783 0.9319 0.0059 -17.750 -0.7639 0.07444 0.07120 -0.0908 0.9114 0.0060 -17.500 -0.7549 0.06647 0.06287 -0.1023 0.8777 0.0060 -17.250 -0.7502 0.06189 0.05801 -0.1073 0.8416 0.0060 -17.000 -0.7532 0.05792 0.05387 -0.1099 0.8281 0.0061 -16.750 -0.7488 0.05509 0.05087 -0.1118 0.8081 0.0062 -16.500 -0.7516 0.05167 0.04729 -0.1136 0.7897 0.0062 -16.250 -0.7499 0.04893 0.04440 -0.1150 0.7753 0.0063 -16.000 -0.7466 0.04650 0.04184 -0.1160 0.7618 0.0063 -15.750 -0.7445 0.04409 0.03928 -0.1169 0.7461 0.0064 -15.500 -0.7433 0.04168 0.03673 -0.1176 0.7338 0.0064 -15.250 -0.7342 0.04012 0.03506 -0.1181 0.7220 0.0065 -15.000 -0.7334 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15.250 1.3522 0.07605 0.07041 -0.0676 0.0065 1.0000 15.500 1.3582 0.07824 0.07263 -0.0674 0.0063 1.0000 15.750 1.3646 0.08039 0.07483 -0.0673 0.0061 1.0000 16.000 1.3695 0.08277 0.07726 -0.0672 0.0058 1.0000 16.250 1.3750 0.08505 0.07958 -0.0671 0.0056 1.0000 16.500 1.3802 0.08737 0.08193 -0.0670 0.0053 1.0000 16.750 1.3836 0.08992 0.08454 -0.0669 0.0051 1.0000 17.000 1.3870 0.09244 0.08709 -0.0669 0.0048 1.0000 17.250 1.3908 0.09490 0.08961 -0.0668 0.0047 1.0000 17.500 1.3933 0.09758 0.09233 -0.0668 0.0046 1.0000 17.750 1.3970 0.10010 0.09490 -0.0669 0.0045 1.0000 18.000 1.4019 0.10248 0.09734 -0.0670 0.0043 1.0000 18.250 1.4041 0.10522 0.10013 -0.0671 0.0042 1.0000 18.500 1.4067 0.10789 0.10286 -0.0673 0.0041 1.0000 18.750 1.4098 0.11047 0.10550 -0.0675 0.0039 1.0000 19.000 1.4117 0.11324 0.10832 -0.0678 0.0038 1.0000 19.250 1.4132 0.11609 0.11122 -0.0681 0.0037 1.0000