XFOIL Version 6.94 Calculated polar for: GOE 416A AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.080 Re = 0.450 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- 0.000 0.1038 0.00906 0.00272 -0.0278 0.6533 0.5343 0.500 0.1610 0.00924 0.00275 -0.0278 0.6169 0.5423 1.000 0.2184 0.00929 0.00275 -0.0280 0.5838 0.5470 1.500 0.2757 0.00941 0.00278 -0.0281 0.5475 0.5527 2.000 0.3329 0.00952 0.00281 -0.0283 0.5105 0.5581 2.500 0.3900 0.00967 0.00291 -0.0284 0.4773 0.5629 3.000 0.4469 0.00992 0.00306 -0.0284 0.4460 0.5691 3.500 0.5033 0.01012 0.00327 -0.0285 0.4167 0.5746 4.500 0.6154 0.01066 0.00375 -0.0285 0.3361 0.5880 5.000 0.6688 0.01130 0.00409 -0.0283 0.2602 0.5957 5.500 0.7225 0.01191 0.00461 -0.0280 0.2249 0.6054 6.000 0.7765 0.01239 0.00515 -0.0278 0.2016 0.6158 6.500 0.8304 0.01291 0.00569 -0.0275 0.1657 0.6294 7.000 0.8798 0.01394 0.00650 -0.0269 0.1018 0.6481 7.500 0.9272 0.01515 0.00767 -0.0259 0.0637 0.6809 8.000 0.9692 0.01578 0.00889 -0.0235 0.0515 0.8458 8.500 1.0163 0.01690 0.01019 -0.0222 0.0459 1.0000 9.000 1.0603 0.01840 0.01174 -0.0207 0.0413 1.0000 9.500 1.0998 0.02018 0.01355 -0.0188 0.0388 1.0000 10.000 1.1413 0.02163 0.01513 -0.0171 0.0365 1.0000 10.500 1.1746 0.02374 0.01724 -0.0146 0.0348 1.0000 11.000 1.2098 0.02544 0.01916 -0.0123 0.0332 1.0000 11.500 1.2356 0.02758 0.02139 -0.0089 0.0316 1.0000 12.000 1.2572 0.02965 0.02365 -0.0058 0.0293 1.0000 12.500 1.2750 0.03258 0.02677 -0.0031 0.0283 1.0000 13.000 1.2917 0.03579 0.03004 -0.0010 0.0271 1.0000 13.500 1.2996 0.03993 0.03451 0.0000 0.0256 1.0000 14.000 1.3079 0.04440 0.03920 -0.0003 0.0244 1.0000 14.500 1.3145 0.04943 0.04436 -0.0009 0.0236 1.0000 15.000 1.3103 0.05611 0.05130 -0.0023 0.0228 1.0000 15.500 1.2986 0.06434 0.05989 -0.0058 0.0220 1.0000 16.000 1.2841 0.07326 0.06910 -0.0095 0.0214 1.0000 16.500 1.2705 0.08269 0.07876 -0.0142 0.0208 1.0000 17.000 1.2612 0.09156 0.08774 -0.0185 0.0202 1.0000 17.500 1.2381 0.10310 0.09950 -0.0241 0.0197 1.0000 18.000 1.1991 0.11886 0.11564 -0.0333 0.0194 1.0000 18.500 1.1510 0.13767 0.13484 -0.0445 0.0193 1.0000