XFOIL Version 6.94 Calculated polar for: GOE 393 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.080 Re = 0.300 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- 0.000 0.3191 0.01098 0.00283 -0.0511 0.5990 0.0569 0.500 0.3756 0.01126 0.00307 -0.0507 0.5808 0.1033 1.000 0.4310 0.01131 0.00305 -0.0503 0.5634 0.1182 1.500 0.4864 0.01126 0.00303 -0.0500 0.5444 0.1322 2.000 0.5416 0.01121 0.00303 -0.0497 0.5255 0.1504 2.500 0.5968 0.01114 0.00304 -0.0493 0.5062 0.1654 3.000 0.6519 0.01105 0.00306 -0.0490 0.4838 0.1953 3.500 0.7074 0.00946 0.00300 -0.0488 0.4484 1.0000 4.000 0.7607 0.00975 0.00294 -0.0481 0.3815 1.0000 4.500 0.8133 0.01040 0.00325 -0.0476 0.3126 1.0000 5.000 0.8660 0.01110 0.00367 -0.0472 0.2485 1.0000 5.500 0.9073 0.01438 0.00548 -0.0465 0.0104 1.0000 6.000 0.9592 0.01522 0.00628 -0.0458 0.0061 1.0000 6.500 1.0109 0.01601 0.00726 -0.0451 0.0057 1.0000 7.000 1.0616 0.01696 0.00852 -0.0442 0.0057 1.0000 7.500 1.1111 0.01808 0.00991 -0.0432 0.0059 1.0000 8.000 1.1582 0.01954 0.01165 -0.0418 0.0062 1.0000 8.500 1.2009 0.02156 0.01400 -0.0402 0.0065 1.0000 9.000 1.2356 0.02441 0.01721 -0.0378 0.0070 1.0000 9.500 1.2608 0.02813 0.02126 -0.0346 0.0075 1.0000 10.000 1.2767 0.03287 0.02643 -0.0308 0.0079 1.0000 10.500 1.2878 0.03781 0.03175 -0.0267 0.0084 1.0000 11.000 1.3026 0.04139 0.03561 -0.0241 0.0091 1.0000 11.500 1.3008 0.04846 0.04329 -0.0219 0.0104 1.0000 12.000 1.2782 0.05861 0.05407 -0.0228 0.0114 1.0000 12.500 1.2482 0.06994 0.06588 -0.0270 0.0119 1.0000 13.000 1.2141 0.08270 0.07904 -0.0334 0.0122 1.0000 13.500 1.1788 0.09677 0.09344 -0.0414 0.0121 1.0000 14.000 1.1444 0.11256 0.10950 -0.0511 0.0119 1.0000 14.500 1.1109 0.13067 0.12785 -0.0626 0.0115 1.0000 15.000 1.0783 0.15169 0.14903 -0.0752 0.0108 1.0000 15.500 1.0417 0.17831 0.17555 -0.0886 0.0109 1.0000 16.000 1.0355 0.19456 0.19174 -0.0967 0.0125 1.0000