XFOIL Version 6.94 Calculated polar for: GOE 118 (MVA MK.7) 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 0.3616 0.01066 0.00291 -0.0777 0.5871 0.1440 0.500 0.4181 0.01052 0.00277 -0.0777 0.5695 0.1491 1.000 0.4744 0.01040 0.00264 -0.0776 0.5538 0.1557 1.500 0.5311 0.01030 0.00262 -0.0776 0.5392 0.1682 2.000 0.5870 0.01007 0.00261 -0.0776 0.5219 0.2225 2.500 0.6445 0.00828 0.00287 -0.0780 0.5091 1.0000 3.000 0.6996 0.00850 0.00299 -0.0776 0.4896 1.0000 3.500 0.7545 0.00877 0.00317 -0.0772 0.4741 1.0000 4.000 0.8096 0.00904 0.00341 -0.0769 0.4571 1.0000 4.500 0.8640 0.00934 0.00365 -0.0765 0.4370 1.0000 5.000 0.9178 0.00967 0.00394 -0.0761 0.4128 1.0000 5.500 0.9710 0.01008 0.00428 -0.0755 0.3857 1.0000 6.000 1.0229 0.01059 0.00470 -0.0749 0.3531 1.0000 6.500 1.0734 0.01124 0.00522 -0.0740 0.3162 1.0000 7.000 1.1216 0.01210 0.00592 -0.0729 0.2722 1.0000 7.500 1.1485 0.01495 0.00785 -0.0693 0.1134 1.0000 8.000 1.1759 0.01752 0.00990 -0.0656 0.0263 1.0000 8.500 1.2133 0.01897 0.01142 -0.0630 0.0216 1.0000 9.000 1.2422 0.02083 0.01339 -0.0594 0.0189 1.0000 9.500 1.2594 0.02296 0.01565 -0.0545 0.0174 1.0000 10.000 1.2649 0.02629 0.01914 -0.0499 0.0167 1.0000 10.500 1.2773 0.02976 0.02281 -0.0474 0.0156 1.0000 11.000 1.2825 0.03439 0.02759 -0.0457 0.0148 1.0000 11.500 1.2759 0.04039 0.03369 -0.0442 0.0138 1.0000 12.000 1.2854 0.04499 0.03848 -0.0432 0.0131 1.0000 12.500 1.2903 0.04997 0.04361 -0.0420 0.0123 1.0000 13.000 1.2962 0.05476 0.04851 -0.0405 0.0119 1.0000 13.500 1.3077 0.05829 0.05201 -0.0367 0.0111 1.0000 14.000 1.3137 0.06361 0.05763 -0.0363 0.0108 1.0000 14.500 1.3205 0.06881 0.06310 -0.0352 0.0105 1.0000 15.000 1.3243 0.07463 0.06920 -0.0345 0.0103 1.0000 15.500 1.3198 0.08182 0.07673 -0.0347 0.0099 1.0000 16.000 1.3098 0.09019 0.08546 -0.0358 0.0099 1.0000 16.500 1.2997 0.09873 0.09431 -0.0374 0.0105 1.0000 17.000 1.2641 0.11242 0.10848 -0.0427 0.0103 1.0000 17.500 1.2323 0.12623 0.12266 -0.0489 0.0106 1.0000 18.000 1.2177 0.13677 0.13337 -0.0535 0.0110 1.0000