XFOIL Version 6.94 Calculated polar for: GOE 113 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 ------- -------- --------- --------- -------- ------- ------- 0.000 0.2158 0.00927 0.00181 -0.0382 0.6583 0.0602 0.500 0.2569 0.00675 0.00176 -0.0357 0.6208 0.8401 1.000 0.3344 0.00677 0.00175 -0.0395 0.5815 1.0000 1.500 0.3869 0.00696 0.00174 -0.0386 0.5509 1.0000 2.000 0.4403 0.00718 0.00182 -0.0379 0.5250 1.0000 2.500 0.4938 0.00742 0.00191 -0.0373 0.4934 1.0000 3.000 0.5473 0.00773 0.00202 -0.0368 0.4445 1.0000 3.500 0.6016 0.00803 0.00223 -0.0363 0.4083 1.0000 4.000 0.6526 0.00881 0.00255 -0.0356 0.3023 1.0000 4.500 0.7034 0.00975 0.00309 -0.0350 0.2301 1.0000 5.000 0.7559 0.01043 0.00365 -0.0345 0.2001 1.0000 5.500 0.8087 0.01104 0.00419 -0.0340 0.1781 1.0000 6.000 0.8614 0.01159 0.00468 -0.0336 0.1579 1.0000 6.500 0.9138 0.01222 0.00534 -0.0330 0.1494 1.0000 7.000 0.9663 0.01274 0.00594 -0.0326 0.1376 1.0000 7.500 1.0182 0.01335 0.00659 -0.0320 0.1271 1.0000 8.000 1.0698 0.01395 0.00727 -0.0315 0.1103 1.0000 8.500 1.1197 0.01480 0.00803 -0.0307 0.0678 1.0000 9.000 1.1617 0.01663 0.00971 -0.0291 0.0486 1.0000 9.500 1.2042 0.01826 0.01148 -0.0274 0.0405 1.0000 10.000 1.2480 0.01957 0.01288 -0.0261 0.0315 1.0000 10.500 1.2894 0.02105 0.01455 -0.0244 0.0253 1.0000 11.000 1.3293 0.02258 0.01618 -0.0225 0.0184 1.0000 11.500 1.3594 0.02494 0.01870 -0.0197 0.0118 1.0000 12.000 1.3793 0.02781 0.02181 -0.0158 0.0093 1.0000 12.500 1.3841 0.03100 0.02525 -0.0107 0.0086 1.0000 13.000 1.3721 0.03631 0.03083 -0.0072 0.0081 1.0000 13.500 1.3655 0.04285 0.03774 -0.0079 0.0075 1.0000 14.000 1.3515 0.05168 0.04688 -0.0114 0.0072 1.0000 14.500 1.3313 0.06200 0.05750 -0.0160 0.0071 1.0000 15.000 1.3046 0.07363 0.06942 -0.0212 0.0071 1.0000 15.500 1.2728 0.08661 0.08264 -0.0272 0.0069 1.0000 16.000 1.2426 0.10016 0.09643 -0.0337 0.0068 1.0000 16.500 1.2162 0.11360 0.11008 -0.0402 0.0067 1.0000 17.000 1.1875 0.12754 0.12429 -0.0466 0.0070 1.0000 17.500 1.1675 0.14063 0.13757 -0.0534 0.0069 1.0000