XFOIL Version 6.96 Calculated polar for: FX 61-168 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.500 e 6 Ncrit = 5.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -13.000 -0.5796 0.07067 0.06810 -0.0658 1.0000 0.0052 -12.750 -0.6342 0.05890 0.05601 -0.0719 1.0000 0.0050 -12.500 -0.6618 0.05287 0.04977 -0.0737 1.0000 0.0051 -12.250 -0.6804 0.04724 0.04387 -0.0762 0.9992 0.0051 -11.750 -0.6867 0.03808 0.03410 -0.0819 0.9903 0.0054 -11.500 -0.6860 0.03494 0.03072 -0.0825 0.9812 0.0054 -11.000 -0.6650 0.03016 0.02548 -0.0840 0.9410 0.0057 -10.750 -0.6128 0.02737 0.02233 -0.0926 0.9193 0.0059 -10.500 -0.5366 0.02456 0.01914 -0.1059 0.8921 0.0063 -10.000 -0.4470 0.02260 0.01660 -0.1169 0.8021 0.0070 -9.750 -0.4301 0.02197 0.01577 -0.1163 0.7736 0.0072 -9.250 -0.3997 0.02065 0.01411 -0.1139 0.7337 0.0079 -9.000 -0.3855 0.01997 0.01328 -0.1125 0.7189 0.0081 -8.750 -0.3715 0.01933 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1.2115 0.01197 0.00566 -0.1173 0.3930 0.6939 6.500 1.2351 0.01224 0.00592 -0.1168 0.3806 0.6957 6.750 1.2577 0.01256 0.00620 -0.1161 0.3661 0.6974 7.000 1.2792 0.01289 0.00653 -0.1152 0.3511 0.6990 7.250 1.3002 0.01325 0.00688 -0.1142 0.3359 0.7006 7.500 1.3181 0.01375 0.00730 -0.1127 0.3138 0.7021 7.750 1.3336 0.01434 0.00779 -0.1108 0.2868 0.7038 8.000 1.3455 0.01500 0.00833 -0.1083 0.2586 0.7057 8.250 1.3527 0.01581 0.00898 -0.1050 0.2278 0.7077 8.500 1.3582 0.01675 0.00976 -0.1016 0.1961 0.7095 8.750 1.3629 0.01774 0.01060 -0.0982 0.1695 0.7112 9.000 1.3675 0.01875 0.01150 -0.0950 0.1446 0.7127 9.250 1.3714 0.01981 0.01249 -0.0918 0.1255 0.7144 9.500 1.3760 0.02090 0.01352 -0.0889 0.1080 0.7161 9.750 1.3789 0.02213 0.01471 -0.0860 0.0923 0.7180 10.000 1.3830 0.02340 0.01597 -0.0835 0.0798 0.7198 10.250 1.3848 0.02492 0.01745 -0.0811 0.0664 0.7215 10.500 1.3866 0.02659 0.01911 -0.0790 0.0561 0.7233 10.750 1.3894 0.02831 0.02084 -0.0773 0.0471 0.7251 11.000 1.3911 0.03025 0.02278 -0.0757 0.0386 0.7267 11.250 1.3923 0.03236 0.02490 -0.0744 0.0315 0.7284 11.500 1.3936 0.03455 0.02712 -0.0733 0.0259 0.7302 11.750 1.3991 0.03648 0.02911 -0.0725 0.0227 0.7321 12.000 1.4021 0.03869 0.03138 -0.0717 0.0194 0.7339 12.250 1.4058 0.04090 0.03365 -0.0711 0.0163 0.7358 12.500 1.4086 0.04326 0.03606 -0.0706 0.0138 0.7376 12.750 1.4110 0.04573 0.03858 -0.0702 0.0113 0.7394 13.000 1.4110 0.04851 0.04141 -0.0698 0.0074 0.7411 13.500 1.4089 0.05455 0.04757 -0.0694 0.0036 0.7443 13.750 1.4105 0.05740 0.05052 -0.0694 0.0033 0.7460 14.000 1.4110 0.06046 0.05367 -0.0695 0.0030 0.7478 14.250 1.4128 0.06345 0.05677 -0.0697 0.0028 0.7500 14.500 1.4150 0.06647 0.05989 -0.0701 0.0027 0.7522 14.750 1.4164 0.06966 0.06319 -0.0705 0.0026 0.7544 15.000 1.4174 0.07298 0.06662 -0.0711 0.0025 0.7566 15.250 1.4182 0.07641 0.07015 -0.0717 0.0024 0.7586 15.500 1.4185 0.07998 0.07384 -0.0725 0.0023 0.7606 15.750 1.4183 0.08371 0.07771 -0.0735 0.0022 0.7627 16.000 1.4174 0.08759 0.08170 -0.0745 0.0022 0.7650 16.250 1.4160 0.09163 0.08586 -0.0757 0.0021 0.7673 16.500 1.4145 0.09577 0.09012 -0.0770 0.0021 0.7697 16.750 1.4119 0.10015 0.09462 -0.0785 0.0020 0.7720 17.000 1.4086 0.10471 0.09931 -0.0802 0.0020 0.7744 17.250 1.4051 0.10938 0.10411 -0.0820 0.0020 0.7769 17.500 1.4001 0.11437 0.10924 -0.0840 0.0019 0.7796