XFOIL Version 6.96 Calculated polar for: NASA SC(2)-0410 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 ------ -------- --------- --------- -------- -------- -------- -17.750 -1.0016 0.13719 0.13502 0.0099 1.0000 0.0102 -17.500 -1.1342 0.10461 0.10207 -0.0079 1.0000 0.0099 -17.250 -1.1990 0.08701 0.08422 -0.0178 1.0000 0.0098 -17.000 -1.2374 0.07476 0.07175 -0.0254 1.0000 0.0097 -16.750 -1.2636 0.06551 0.06231 -0.0314 1.0000 0.0097 -16.500 -1.2827 0.05843 0.05507 -0.0357 1.0000 0.0097 -16.250 -1.2974 0.05286 0.04934 -0.0386 1.0000 0.0098 -16.000 -1.3080 0.04832 0.04466 -0.0405 1.0000 0.0098 -15.750 -1.3151 0.04451 0.04072 -0.0418 1.0000 0.0099 -15.500 -1.3190 0.04125 0.03734 -0.0425 1.0000 0.0099 -15.250 -1.3201 0.03846 0.03444 -0.0430 1.0000 0.0100 -15.000 -1.3183 0.03605 0.03193 -0.0431 1.0000 0.0101 -14.750 -1.3140 0.03397 0.02975 -0.0431 1.0000 0.0102 -14.500 -1.3080 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0.7983 0.01138 0.00441 -0.0706 0.0277 0.7061 5.250 0.8259 0.01161 0.00465 -0.0705 0.0255 0.7080 5.500 0.8536 0.01183 0.00488 -0.0705 0.0233 0.7097 5.750 0.8810 0.01211 0.00515 -0.0704 0.0211 0.7116 6.000 0.9083 0.01237 0.00542 -0.0703 0.0193 0.7135 6.250 0.9354 0.01268 0.00573 -0.0701 0.0176 0.7155 6.500 0.9625 0.01298 0.00605 -0.0700 0.0168 0.7174 6.750 0.9894 0.01328 0.00638 -0.0698 0.0161 0.7193 7.000 1.0162 0.01361 0.00673 -0.0696 0.0155 0.7211 7.250 1.0427 0.01397 0.00711 -0.0693 0.0149 0.7230 7.500 1.0689 0.01437 0.00755 -0.0690 0.0143 0.7251 7.750 1.0949 0.01479 0.00801 -0.0687 0.0139 0.7273 8.000 1.1211 0.01516 0.00844 -0.0684 0.0137 0.7297 8.250 1.1470 0.01557 0.00890 -0.0681 0.0134 0.7321 8.500 1.1726 0.01599 0.00937 -0.0677 0.0132 0.7345 8.750 1.1980 0.01644 0.00988 -0.0673 0.0130 0.7368 9.250 1.2481 0.01739 0.01095 -0.0664 0.0125 0.7416 9.500 1.2727 0.01790 0.01153 -0.0659 0.0123 0.7445 9.750 1.2970 0.01843 0.01212 -0.0654 0.0121 0.7476 10.000 1.3209 0.01898 0.01273 -0.0648 0.0119 0.7508 10.250 1.3444 0.01959 0.01340 -0.0642 0.0116 0.7538 10.500 1.3670 0.02030 0.01418 -0.0634 0.0114 0.7566 10.750 1.3884 0.02113 0.01512 -0.0625 0.0111 0.7597 11.000 1.4090 0.02203 0.01612 -0.0615 0.0110 0.7630 11.250 1.4308 0.02272 0.01690 -0.0607 0.0109 0.7664 11.500 1.4519 0.02345 0.01774 -0.0597 0.0108 0.7697 11.750 1.4724 0.02422 0.01861 -0.0587 0.0107 0.7728 12.000 1.4921 0.02504 0.01954 -0.0576 0.0106 0.7762 12.250 1.5108 0.02591 0.02053 -0.0564 0.0105 0.7799 12.500 1.5286 0.02684 0.02157 -0.0551 0.0104 0.7836 13.000 1.5604 0.02889 0.02387 -0.0520 0.0103 0.7909 13.250 1.5741 0.03000 0.02511 -0.0503 0.0102 0.7951 13.500 1.5859 0.03120 0.02643 -0.0483 0.0101 0.7994 13.750 1.5952 0.03246 0.02783 -0.0460 0.0100 0.8033 14.000 1.5999 0.03375 0.02926 -0.0430 0.0099 0.8077 14.250 1.5994 0.03520 0.03084 -0.0394 0.0099 0.8124 14.750 1.5978 0.03877 0.03470 -0.0339 0.0097 0.8214 15.000 1.5939 0.04107 0.03716 -0.0316 0.0097 0.8263 15.250 1.5882 0.04377 0.04001 -0.0300 0.0096 0.8311 15.500 1.5810 0.04691 0.04331 -0.0291 0.0096 0.8362 15.750 1.5695 0.05092 0.04750 -0.0290 0.0095 0.8417 16.000 1.5540 0.05603 0.05279 -0.0303 0.0095 0.8466 16.250 1.5316 0.06304 0.06002 -0.0337 0.0095 0.8511 16.500 1.4949 0.07446 0.07170 -0.0414 0.0095 0.8548 16.750 1.3409 0.11385 0.11162 -0.0663 0.0098 0.8484