XFOIL Version 6.94 Calculated polar for: GOE 375 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 0.4301 0.01775 0.01177 -0.0717 0.7019 0.0358 1.500 0.5737 0.00347 -0.00443 -0.0608 0.4945 0.0740 2.000 0.6230 0.00344 -0.00470 -0.0593 0.4449 0.0634 2.500 0.6720 0.00344 -0.00489 -0.0580 0.4010 0.0579 3.000 0.7211 0.00349 -0.00500 -0.0569 0.3591 0.0568 3.500 0.7701 0.00364 -0.00505 -0.0559 0.3216 0.0604 4.000 0.8186 0.00389 -0.00500 -0.0549 0.2935 0.0683 4.500 0.8760 0.00342 -0.00396 -0.0562 0.2710 1.0000 5.000 0.9231 0.00366 -0.00404 -0.0552 0.2225 1.0000 5.500 0.9695 0.00403 -0.00396 -0.0540 0.1809 1.0000 6.000 1.0140 0.00460 -0.00348 -0.0526 0.1308 1.0000 6.500 1.0479 0.00616 -0.00253 -0.0499 0.0307 1.0000 7.000 1.0897 0.00712 -0.00152 -0.0478 0.0074 1.0000 7.500 1.1334 0.00787 -0.00056 -0.0459 0.0084 1.0000 8.000 1.1752 0.00880 0.00067 -0.0436 0.0101 1.0000 8.500 1.2123 0.01017 0.00238 -0.0406 0.0110 1.0000 9.000 1.2378 0.01231 0.00494 -0.0363 0.0116 1.0000 9.500 1.2413 0.01510 0.00800 -0.0295 0.0119 1.0000 10.000 1.2237 0.01836 0.01145 -0.0212 0.0121 1.0000 10.500 1.2195 0.02182 0.01519 -0.0164 0.0127 1.0000 11.000 1.2075 0.02746 0.02105 -0.0131 0.0134 1.0000 11.500 1.2258 0.03344 0.02689 -0.0098 0.0146 1.0000 12.000 1.3923 0.04630 0.03950 -0.0113 0.0170 1.0000 12.500 1.3741 0.05088 0.04461 -0.0061 0.0166 1.0000 13.000 1.3573 0.05700 0.05120 -0.0036 0.0162 1.0000 13.500 1.3352 0.06444 0.05908 -0.0034 0.0159 1.0000 14.000 1.3101 0.07340 0.06845 -0.0056 0.0158 1.0000 14.500 1.2815 0.08406 0.07952 -0.0101 0.0162 1.0000 15.000 1.2572 0.09558 0.09136 -0.0155 0.0166 1.0000 15.500 1.2327 0.10800 0.10406 -0.0222 0.0169 1.0000 16.000 1.2140 0.12048 0.11677 -0.0288 0.0170 1.0000