XFOIL Version 6.94 Calculated polar for: GOE 655 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 ------- -------- --------- --------- -------- ------- ------- 0.500 0.5439 0.00831 0.00305 -0.0881 0.4939 0.9574 1.000 0.6345 0.00883 0.00324 -0.0954 0.4519 0.9961 1.500 0.6886 0.00923 0.00337 -0.0952 0.4227 1.0000 2.000 0.7293 0.00957 0.00354 -0.0921 0.4020 1.0000 2.500 0.7714 0.00992 0.00374 -0.0893 0.3867 1.0000 3.000 0.8136 0.01032 0.00398 -0.0866 0.3735 1.0000 3.500 0.8584 0.01063 0.00427 -0.0843 0.3635 1.0000 4.000 0.8975 0.01104 0.00445 -0.0810 0.3319 1.0000 4.500 0.9401 0.01140 0.00472 -0.0784 0.3161 1.0000 5.000 0.9783 0.01190 0.00501 -0.0751 0.2862 1.0000 5.500 1.0145 0.01251 0.00537 -0.0715 0.2469 1.0000 6.000 1.0383 0.01356 0.00602 -0.0658 0.1850 1.0000 6.500 1.0604 0.01469 0.00690 -0.0600 0.1495 1.0000 7.000 1.0858 0.01578 0.00779 -0.0550 0.1178 1.0000 7.500 1.0755 0.01874 0.01021 -0.0452 0.0039 1.0000 8.000 1.1072 0.01971 0.01123 -0.0418 0.0034 1.0000 8.500 1.1375 0.02082 0.01241 -0.0386 0.0034 1.0000 9.000 1.1660 0.02211 0.01380 -0.0354 0.0034 1.0000 9.500 1.1921 0.02364 0.01545 -0.0322 0.0035 1.0000 10.000 1.2156 0.02548 0.01742 -0.0292 0.0036 1.0000 10.500 1.2358 0.02772 0.01982 -0.0265 0.0037 1.0000 11.000 1.2518 0.03050 0.02278 -0.0239 0.0038 1.0000 11.500 1.2615 0.03409 0.02658 -0.0218 0.0039 1.0000 12.000 1.2703 0.03806 0.03073 -0.0202 0.0041 1.0000 12.500 1.2773 0.04250 0.03535 -0.0192 0.0043 1.0000 13.000 1.2776 0.04797 0.04103 -0.0188 0.0045 1.0000 13.500 1.2700 0.05476 0.04804 -0.0192 0.0046 1.0000 14.000 1.2553 0.06293 0.05646 -0.0204 0.0048 1.0000 14.500 1.2340 0.07243 0.06621 -0.0223 0.0050 1.0000 15.000 1.2085 0.08291 0.07692 -0.0250 0.0051 1.0000 15.500 1.1827 0.09373 0.08796 -0.0279 0.0052 1.0000 16.000 1.1577 0.10464 0.09906 -0.0312 0.0052 1.0000 16.500 1.1348 0.11548 0.11006 -0.0347 0.0053 1.0000 17.000 1.1344 0.12288 0.11760 -0.0370 0.0056 1.0000 17.500 1.1294 0.13111 0.12598 -0.0400 0.0059 1.0000 18.000 1.1257 0.13913 0.13412 -0.0430 0.0063 1.0000 18.500 1.1314 0.14480 0.13978 -0.0449 0.0068 1.0000 19.000 1.1416 0.15060 0.14573 -0.0473 0.0075 1.0000 19.500 1.1740 0.15054 0.14558 -0.0459 0.0088 1.0000 20.000 1.2094 0.15053 0.14565 -0.0443 0.0110 1.0000 20.500 1.0464 0.15645 0.15268 -0.0611 0.0109 1.0000