XFOIL Version 6.96 Calculated polar for: FX 61-163 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 = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -13.500 -0.4123 0.13682 0.13513 -0.0297 1.0001 0.0086 -13.250 -0.4106 0.13235 0.13067 -0.0315 1.0001 0.0090 -13.000 -0.4106 0.12715 0.12549 -0.0335 1.0001 0.0094 -12.750 -0.4125 0.11959 0.11795 -0.0367 1.0001 0.0101 -12.500 -0.4144 0.11366 0.11203 -0.0392 1.0001 0.0101 -12.250 -0.4186 0.10709 0.10548 -0.0419 1.0001 0.0102 -11.750 -0.4550 0.08312 0.08143 -0.0547 1.0001 0.0102 -11.500 -0.4770 0.07490 0.07312 -0.0603 1.0001 0.0102 -11.250 -0.4979 0.06858 0.06670 -0.0640 1.0001 0.0102 -11.000 -0.5185 0.06342 0.06145 -0.0662 1.0001 0.0102 -10.750 -0.5386 0.05906 0.05699 -0.0673 1.0001 0.0102 -10.500 -0.5481 0.05363 0.05138 -0.0722 0.9901 0.0102 -10.250 -0.5400 0.04703 0.04444 -0.0829 0.9486 0.0102 -9.000 -0.4915 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5.750 1.1632 0.00913 0.00397 -0.1168 0.4144 0.6968 6.000 1.1886 0.00945 0.00420 -0.1166 0.3933 0.6985 6.250 1.2136 0.00979 0.00445 -0.1164 0.3687 0.7001 6.500 1.2385 0.01015 0.00472 -0.1161 0.3489 0.7014 6.750 1.2635 0.01045 0.00498 -0.1159 0.3289 0.7034 7.000 1.2867 0.01089 0.00532 -0.1154 0.3024 0.7051 7.250 1.3072 0.01153 0.00578 -0.1144 0.2656 0.7068 7.500 1.3258 0.01229 0.00632 -0.1132 0.2250 0.7085 7.750 1.3431 0.01312 0.00694 -0.1118 0.1859 0.7102 8.000 1.3603 0.01391 0.00756 -0.1103 0.1540 0.7119 8.250 1.3769 0.01470 0.00819 -0.1088 0.1265 0.7136 8.500 1.3932 0.01546 0.00882 -0.1073 0.1025 0.7151 9.250 1.4322 0.01787 0.01096 -0.1010 0.0492 0.7198 9.500 1.4417 0.01863 0.01169 -0.0983 0.0378 0.7215 9.750 1.4484 0.01964 0.01265 -0.0954 0.0259 0.7234 10.000 1.4578 0.02055 0.01355 -0.0931 0.0201 0.7251 10.250 1.4658 0.02159 0.01460 -0.0907 0.0151 0.7268 10.500 1.4726 0.02278 0.01580 -0.0885 0.0107 0.7285 10.750 1.4764 0.02430 0.01732 -0.0863 0.0057 0.7300 11.000 1.4823 0.02582 0.01888 -0.0847 0.0042 0.7315 11.250 1.4903 0.02733 0.02047 -0.0835 0.0038 0.7335 11.500 1.4974 0.02901 0.02225 -0.0825 0.0035 0.7355 11.750 1.5037 0.03087 0.02419 -0.0816 0.0033 0.7374 12.000 1.5089 0.03289 0.02630 -0.0808 0.0030 0.7392 12.250 1.5131 0.03506 0.02855 -0.0801 0.0029 0.7410 12.500 1.5162 0.03739 0.03098 -0.0794 0.0028 0.7428 12.750 1.5200 0.03965 0.03332 -0.0788 0.0027 0.7445 13.000 1.5228 0.04203 0.03578 -0.0783 0.0027 0.7460 13.250 1.5244 0.04457 0.03842 -0.0778 0.0026 0.7481 13.500 1.5244 0.04728 0.04124 -0.0772 0.0026 0.7502 13.750 1.5239 0.05005 0.04410 -0.0767 0.0026 0.7523 14.000 1.5231 0.05291 0.04705 -0.0763 0.0025 0.7544 14.250 1.5212 0.05596 0.05020 -0.0760 0.0025 0.7565 14.500 1.5184 0.05916 0.05351 -0.0758 0.0025 0.7584 14.750 1.5161 0.06239 0.05683 -0.0756 0.0024 0.7602 15.000 1.5133 0.06577 0.06031 -0.0756 0.0024 0.7624 15.250 1.5093 0.06937 0.06402 -0.0758 0.0024 0.7647 15.500 1.5045 0.07311 0.06787 -0.0760 0.0023 0.7669 15.750 1.4993 0.07700 0.07186 -0.0763 0.0023 0.7693 16.000 1.4939 0.08099 0.07596 -0.0768 0.0023 0.7718 16.250 1.4876 0.08518 0.08024 -0.0773 0.0023 0.7742 16.500 1.4814 0.08938 0.08455 -0.0780 0.0023 0.7767 16.750 1.4746 0.09377 0.08906 -0.0789 0.0023 0.7795 17.000 1.4674 0.09828 0.09368 -0.0798 0.0022 0.7826 17.250 1.4587 0.10312 0.09863 -0.0810 0.0022 0.7860 17.500 1.4499 0.10797 0.10360 -0.0823 0.0022 0.7894 17.750 1.4414 0.11286 0.10861 -0.0837 0.0022 0.7932 18.000 1.4338 0.11770 0.11356 -0.0853 0.0022 0.7977 18.250 1.4251 0.12281 0.11878 -0.0871 0.0022 0.8026 18.500 1.4157 0.12810 0.12420 -0.0891 0.0022 0.8077 18.750 1.4083 0.13307 0.12928 -0.0912 0.0021 0.8145 19.000 1.4003 0.13825 0.13458 -0.0934 0.0021 0.8223