XFOIL Version 6.96 Calculated polar for: FX 83-W-108 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.100 e 6 Ncrit = 5.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -9.500 -0.3862 0.09814 0.09307 -0.0472 1.0000 0.0367 -9.250 -0.3966 0.09567 0.09068 -0.0457 1.0000 0.0365 -9.000 -0.4079 0.09305 0.08815 -0.0445 0.9996 0.0363 -8.750 -0.4003 0.08714 0.08225 -0.0505 0.9938 0.0362 -8.500 -0.3977 0.08039 0.07551 -0.0574 0.9865 0.0364 -8.250 -0.4013 0.07260 0.06774 -0.0655 0.9762 0.0365 -8.000 -0.4091 0.05745 0.05240 -0.0853 0.9610 0.0361 -7.750 -0.4061 0.04796 0.04237 -0.0971 0.9499 0.0361 -7.500 -0.3897 0.04116 0.03484 -0.1041 0.9431 0.0368 -7.250 -0.3722 0.03704 0.03012 -0.1065 0.9359 0.0372 -7.000 -0.3473 0.03352 0.02591 -0.1089 0.9310 0.0379 -6.750 -0.3176 0.03104 0.02320 -0.1110 0.9281 0.0390 -6.500 -0.2941 0.02956 0.02151 -0.1114 0.9224 0.0403 -6.250 -0.2665 0.02827 0.02000 -0.1123 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0.01809 0.01030 -0.1179 0.7390 1.0000 3.750 0.8582 0.01813 0.01039 -0.1171 0.7297 1.0000 4.000 0.8827 0.01815 0.01047 -0.1161 0.7194 1.0000 4.250 0.9134 0.01790 0.01025 -0.1159 0.7088 1.0000 4.500 0.9412 0.01766 0.01004 -0.1152 0.6950 1.0000 4.750 0.9621 0.01760 0.01003 -0.1133 0.6779 1.0000 5.000 0.9823 0.01758 0.01007 -0.1114 0.6595 1.0000 5.250 1.0040 0.01757 0.01011 -0.1098 0.6403 1.0000 5.500 1.0280 0.01751 0.01006 -0.1085 0.6193 1.0000 5.750 1.0458 0.01762 0.01021 -0.1062 0.5917 1.0000 6.000 1.0656 0.01771 0.01025 -0.1041 0.5565 1.0000 6.250 1.0837 0.01794 0.01032 -0.1018 0.5094 1.0000 6.500 1.0963 0.01850 0.01050 -0.0987 0.4411 1.0000 6.750 1.0999 0.01949 0.01104 -0.0943 0.3661 1.0000 7.000 1.0999 0.02072 0.01182 -0.0897 0.3009 1.0000 7.250 1.1021 0.02201 0.01276 -0.0858 0.2461 1.0000 7.500 1.1070 0.02327 0.01375 -0.0825 0.2023 1.0000 7.750 1.1141 0.02448 0.01476 -0.0796 0.1684 1.0000 8.000 1.1222 0.02569 0.01581 -0.0770 0.1418 1.0000 8.250 1.1311 0.02689 0.01693 -0.0746 0.1227 1.0000 8.500 1.1404 0.02810 0.01809 -0.0723 0.1084 1.0000 8.750 1.1502 0.02931 0.01929 -0.0702 0.0974 1.0000 9.000 1.1590 0.03062 0.02058 -0.0680 0.0884 1.0000 9.250 1.1688 0.03191 0.02188 -0.0661 0.0809 1.0000 9.500 1.1782 0.03327 0.02330 -0.0641 0.0750 1.0000 9.750 1.1886 0.03459 0.02467 -0.0623 0.0695 1.0000 10.000 1.1971 0.03611 0.02618 -0.0605 0.0652 1.0000 10.250 1.2088 0.03745 0.02765 -0.0589 0.0610 1.0000 10.500 1.2193 0.03887 0.02912 -0.0574 0.0575 1.0000 10.750 1.2279 0.04058 0.03080 -0.0558 0.0548 1.0000 11.000 1.2412 0.04199 0.03241 -0.0545 0.0519 1.0000 11.250 1.2540 0.04355 0.03409 -0.0531 0.0494 1.0000 11.500 1.2659 0.04515 0.03578 -0.0519 0.0474 1.0000 11.750 1.2768 0.04685 0.03751 -0.0507 0.0455 1.0000 12.000 1.2894 0.04869 0.03945 -0.0495 0.0436 1.0000 12.250 1.3014 0.05052 0.04152 -0.0484 0.0419 1.0000 12.500 1.3117 0.05250 0.04372 -0.0472 0.0402 1.0000 12.750 1.3217 0.05458 0.04597 -0.0461 0.0390 1.0000 13.000 1.3297 0.05677 0.04830 -0.0450 0.0379 1.0000 13.250 1.3371 0.05906 0.05072 -0.0440 0.0371 1.0000 13.500 1.3442 0.06152 0.05326 -0.0430 0.0362 1.0000 13.750 1.3461 0.06457 0.05653 -0.0420 0.0354 1.0000 14.000 1.3414 0.06797 0.06028 -0.0410 0.0347 1.0000 14.250 1.3340 0.07173 0.06436 -0.0402 0.0341 1.0000 14.500 1.3246 0.07574 0.06867 -0.0398 0.0335 1.0000 14.750 1.3125 0.08023 0.07344 -0.0397 0.0330 1.0000 15.000 1.2981 0.08511 0.07860 -0.0402 0.0326 1.0000 15.250 1.2814 0.09060 0.08437 -0.0413 0.0324 1.0000 15.500 1.2613 0.09692 0.09095 -0.0432 0.0322 1.0000 15.750 1.2380 0.10419 0.09849 -0.0462 0.0322 1.0000 16.000 1.2105 0.11288 0.10745 -0.0507 0.0323 1.0000 16.250 1.1753 0.12423 0.11906 -0.0577 0.0326 1.0000 16.500 1.1245 0.14163 0.13671 -0.0694 0.0334 1.0000