XFOIL Version 6.94 Calculated polar for: GOE 342 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 ------- -------- --------- --------- -------- ------- ------- 1.500 0.6914 0.00005 -0.00634 -0.1128 0.7231 0.0648 2.000 0.7426 0.00004 -0.00637 -0.1118 0.6947 0.0656 2.500 0.7932 0.00003 -0.00632 -0.1106 0.6522 0.0688 3.000 0.8400 0.00003 -0.00664 -0.1088 0.5441 0.0766 3.500 0.8816 0.00008 -0.00714 -0.1065 0.4371 0.0963 4.000 0.9237 0.00006 -0.00595 -0.1045 0.3959 1.0000 4.500 0.9705 0.00011 -0.00614 -0.1031 0.3687 1.0000 5.500 1.1469 0.01298 0.00588 -0.1164 0.3295 1.0000 6.000 1.1970 0.01363 0.00651 -0.1152 0.3070 1.0000 6.500 1.2464 0.01426 0.00713 -0.1140 0.2793 1.0000 7.000 1.2947 0.01494 0.00774 -0.1127 0.2271 1.0000 7.500 1.3268 0.01742 0.00940 -0.1094 0.1107 1.0000 8.000 1.3629 0.01940 0.01123 -0.1063 0.0832 1.0000 8.500 1.4014 0.02095 0.01278 -0.1036 0.0679 1.0000 9.000 1.4352 0.02274 0.01458 -0.1002 0.0591 1.0000 9.500 1.4648 0.02471 0.01663 -0.0962 0.0536 1.0000 10.000 1.4862 0.02694 0.01890 -0.0911 0.0496 1.0000 10.500 1.5081 0.02921 0.02131 -0.0863 0.0464 1.0000 11.000 1.5290 0.03172 0.02397 -0.0820 0.0435 1.0000 11.500 1.5506 0.03444 0.02675 -0.0780 0.0412 1.0000 12.000 1.5690 0.03726 0.02985 -0.0742 0.0387 1.0000 12.500 1.5915 0.04051 0.03315 -0.0708 0.0368 1.0000 13.000 1.6051 0.04426 0.03726 -0.0671 0.0352 1.0000 13.500 1.6177 0.04812 0.04137 -0.0638 0.0336 1.0000 14.000 1.6267 0.05234 0.04580 -0.0610 0.0321 1.0000 14.500 1.6215 0.05805 0.05199 -0.0582 0.0312 1.0000 15.000 1.6158 0.06406 0.05830 -0.0564 0.0296 1.0000 15.500 1.6045 0.07067 0.06515 -0.0557 0.0281 1.0000 16.000 1.5813 0.08004 0.07500 -0.0556 0.0281 1.0000 17.000 1.5099 0.10502 0.10093 -0.0651 0.0267 1.0000