XFOIL Version 6.96 Calculated polar for: GOE 506 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 ------ -------- --------- --------- -------- -------- -------- -18.500 -0.5932 0.11402 0.11168 -0.0808 0.9939 0.0072 -18.250 -0.6200 0.10478 0.10232 -0.0861 0.9926 0.0072 -18.000 -0.6438 0.09632 0.09375 -0.0913 0.9909 0.0073 -17.750 -0.6598 0.08935 0.08667 -0.0958 0.9887 0.0073 -17.500 -0.6735 0.08280 0.08003 -0.1004 0.9864 0.0074 -17.250 -0.6904 0.07632 0.07345 -0.1047 0.9821 0.0072 -17.000 -0.7064 0.07010 0.06713 -0.1089 0.9766 0.0074 -16.750 -0.7227 0.06365 0.06058 -0.1140 0.9722 0.0075 -16.500 -0.7350 0.05753 0.05435 -0.1192 0.9665 0.0075 -16.250 -0.7454 0.05099 0.04767 -0.1257 0.9620 0.0075 -16.000 -0.7502 0.04497 0.04152 -0.1323 0.9554 0.0076 -15.750 -0.7468 0.03924 0.03562 -0.1398 0.9493 0.0076 -15.500 -0.7355 0.03446 0.03068 -0.1471 0.9425 0.0077 -15.250 -0.7178 0.03010 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0.01212 0.00614 -0.0763 0.2386 1.0000 5.250 0.9336 0.01247 0.00641 -0.0742 0.2235 1.0000 5.500 0.9365 0.01337 0.00703 -0.0702 0.1751 1.0000 5.750 0.9513 0.01381 0.00742 -0.0682 0.1659 1.0000 6.000 0.9687 0.01417 0.00775 -0.0666 0.1606 1.0000 6.250 0.9870 0.01451 0.00806 -0.0653 0.1574 1.0000 6.500 1.0053 0.01485 0.00840 -0.0639 0.1551 1.0000 6.750 1.0236 0.01521 0.00874 -0.0626 0.1529 1.0000 7.000 1.0429 0.01553 0.00907 -0.0615 0.1516 1.0000 7.250 1.0622 0.01587 0.00941 -0.0604 0.1504 1.0000 7.500 1.0802 0.01627 0.00981 -0.0592 0.1470 1.0000 7.750 1.0997 0.01662 0.01017 -0.0582 0.1456 1.0000 8.000 1.1195 0.01695 0.01052 -0.0573 0.1446 1.0000 8.250 1.1388 0.01731 0.01089 -0.0563 0.1431 1.0000 8.500 1.1577 0.01770 0.01129 -0.0553 0.1418 1.0000 8.750 1.1763 0.01812 0.01172 -0.0543 0.1400 1.0000 9.000 1.1947 0.01855 0.01216 -0.0533 0.1383 1.0000 9.250 1.2116 0.01907 0.01268 -0.0521 0.1351 1.0000 9.500 1.2289 0.01958 0.01319 -0.0510 0.1320 1.0000 9.750 1.2481 0.01999 0.01363 -0.0502 0.1305 1.0000 10.000 1.2660 0.02047 0.01412 -0.0492 0.1270 1.0000 10.250 1.2811 0.02113 0.01473 -0.0479 0.1208 1.0000 10.500 1.2821 0.02265 0.01599 -0.0451 0.0812 1.0000 10.750 1.2899 0.02379 0.01711 -0.0430 0.0696 1.0000 11.000 1.2837 0.02586 0.01903 -0.0396 0.0357 1.0000 11.250 1.2854 0.02751 0.02063 -0.0371 0.0176 1.0000 11.500 1.2964 0.02856 0.02170 -0.0358 0.0146 1.0000 11.750 1.3069 0.02966 0.02283 -0.0345 0.0129 1.0000 12.000 1.3184 0.03073 0.02394 -0.0333 0.0119 1.0000 12.250 1.3293 0.03186 0.02511 -0.0322 0.0112 1.0000 12.500 1.3398 0.03305 0.02634 -0.0311 0.0106 1.0000 12.750 1.3485 0.03441 0.02773 -0.0299 0.0098 1.0000 13.000 1.3567 0.03583 0.02920 -0.0287 0.0093 1.0000 13.250 1.3661 0.03719 0.03062 -0.0278 0.0091 1.0000 13.500 1.3740 0.03871 0.03218 -0.0268 0.0086 1.0000 13.750 1.3815 0.04030 0.03383 -0.0258 0.0083 1.0000 14.000 1.3877 0.04204 0.03562 -0.0249 0.0080 1.0000 14.250 1.3921 0.04399 0.03762 -0.0239 0.0075 1.0000 14.500 1.3954 0.04607 0.03976 -0.0231 0.0071 1.0000 14.750 1.3980 0.04830 0.04205 -0.0223 0.0070 1.0000 15.000 1.4021 0.05044 0.04426 -0.0217 0.0068 1.0000 15.250 1.4043 0.05284 0.04673 -0.0211 0.0065 1.0000 15.500 1.4045 0.05552 0.04950 -0.0207 0.0064 1.0000 15.750 1.4042 0.05833 0.05238 -0.0204 0.0062 1.0000 16.000 1.4023 0.06142 0.05555 -0.0202 0.0061 1.0000 16.250 1.3988 0.06481 0.05903 -0.0203 0.0060 1.0000 16.500 1.3936 0.06852 0.06283 -0.0205 0.0058 1.0000 16.750 1.3890 0.07224 0.06663 -0.0208 0.0057 1.0000 17.000 1.3810 0.07642 0.07091 -0.0213 0.0057 1.0000 17.250 1.3694 0.08116 0.07574 -0.0220 0.0053 1.0000 17.500 1.3612 0.08552 0.08021 -0.0228 0.0053 1.0000 17.750 1.3496 0.09038 0.08517 -0.0237 0.0054 1.0000 18.000 1.3358 0.09559 0.09049 -0.0248 0.0053 1.0000 18.250 1.3250 0.10043 0.09543 -0.0260 0.0051 1.0000