XFOIL Version 6.94 Calculated polar for: GOE 177 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 0.2836 0.01091 0.00283 -0.0555 0.5941 0.1552 0.500 0.3388 0.01079 0.00270 -0.0550 0.5595 0.1725 2.000 0.5034 0.00913 0.00286 -0.0534 0.4589 1.0000 2.500 0.5576 0.00947 0.00298 -0.0527 0.4251 1.0000 3.000 0.6121 0.00986 0.00320 -0.0522 0.4012 1.0000 3.500 0.6663 0.01024 0.00343 -0.0516 0.3756 1.0000 4.000 0.7206 0.01066 0.00376 -0.0511 0.3544 1.0000 4.500 0.7747 0.01109 0.00410 -0.0505 0.3342 1.0000 5.000 0.8291 0.01149 0.00452 -0.0500 0.3169 1.0000 5.500 0.8829 0.01191 0.00492 -0.0495 0.2978 1.0000 6.000 0.9364 0.01235 0.00537 -0.0489 0.2743 1.0000 6.500 0.9898 0.01281 0.00585 -0.0483 0.2484 1.0000 7.000 1.0414 0.01348 0.00640 -0.0476 0.1973 1.0000 7.500 1.0820 0.01583 0.00806 -0.0459 0.0814 1.0000 8.000 1.1284 0.01724 0.00936 -0.0446 0.0545 1.0000 8.500 1.1733 0.01869 0.01090 -0.0431 0.0496 1.0000 9.000 1.2129 0.02062 0.01293 -0.0410 0.0467 1.0000 9.500 1.2489 0.02264 0.01509 -0.0386 0.0447 1.0000 10.000 1.2757 0.02534 0.01785 -0.0352 0.0426 1.0000 10.500 1.3043 0.02772 0.02038 -0.0320 0.0409 1.0000 11.000 1.3283 0.03067 0.02329 -0.0285 0.0389 1.0000 11.500 1.3499 0.03341 0.02627 -0.0248 0.0369 1.0000 12.000 1.3766 0.03683 0.02972 -0.0222 0.0338 1.0000 12.500 1.3876 0.03994 0.03305 -0.0192 0.0308 1.0000 13.000 1.3920 0.04426 0.03774 -0.0170 0.0286 1.0000 13.500 1.4088 0.04819 0.04169 -0.0155 0.0260 1.0000 14.000 1.3911 0.05502 0.04910 -0.0169 0.0244 1.0000 14.500 1.3939 0.06052 0.05473 -0.0179 0.0227 1.0000 15.000 1.3726 0.06979 0.06449 -0.0213 0.0213 1.0000 15.500 1.3582 0.07844 0.07343 -0.0255 0.0199 1.0000 16.000 1.3392 0.08840 0.08366 -0.0304 0.0186 1.0000 16.500 1.3123 0.10093 0.09657 -0.0376 0.0169 1.0000 17.000 1.2909 0.11251 0.10846 -0.0428 0.0168 1.0000 17.500 1.2800 0.12253 0.11862 -0.0481 0.0162 1.0000 18.000 1.2325 0.14198 0.13856 -0.0603 0.0150 1.0000 18.500 1.1719 0.16629 0.16331 -0.0750 0.0153 1.0000