XFOIL Version 6.94 Calculated polar for: GOE 436 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 0.4039 0.00952 0.00225 -0.0764 0.6259 0.1039 0.500 0.4506 0.00929 0.00216 -0.0744 0.5812 0.1838 1.000 0.5358 0.00770 0.00251 -0.0808 0.5157 0.9769 1.500 0.6207 0.00828 0.00264 -0.0872 0.4600 1.0000 2.000 0.6624 0.00868 0.00281 -0.0843 0.4345 1.0000 2.500 0.7064 0.00902 0.00303 -0.0818 0.4165 1.0000 3.000 0.7517 0.00934 0.00327 -0.0795 0.4011 1.0000 3.500 0.7933 0.00975 0.00342 -0.0766 0.3628 1.0000 4.000 0.8358 0.01017 0.00358 -0.0740 0.3213 1.0000 4.500 0.8805 0.01055 0.00385 -0.0717 0.2977 1.0000 5.000 0.9189 0.01132 0.00421 -0.0685 0.2286 1.0000 5.500 0.9462 0.01292 0.00516 -0.0636 0.1217 1.0000 6.000 0.9730 0.01458 0.00635 -0.0587 0.0369 1.0000 6.500 1.0074 0.01563 0.00727 -0.0549 0.0040 1.0000 7.000 1.0465 0.01630 0.00800 -0.0518 0.0037 1.0000 7.500 1.0821 0.01704 0.00882 -0.0482 0.0037 1.0000 8.000 1.1130 0.01784 0.00972 -0.0438 0.0038 1.0000 8.500 1.1420 0.01879 0.01078 -0.0393 0.0039 1.0000 9.000 1.1686 0.01993 0.01205 -0.0347 0.0041 1.0000 9.500 1.1920 0.02131 0.01359 -0.0300 0.0043 1.0000 10.000 1.2107 0.02305 0.01552 -0.0252 0.0045 1.0000 10.500 1.2272 0.02511 0.01774 -0.0208 0.0047 1.0000 11.000 1.2447 0.02729 0.02008 -0.0172 0.0050 1.0000 11.500 1.2561 0.03019 0.02315 -0.0138 0.0053 1.0000 12.000 1.2584 0.03420 0.02737 -0.0108 0.0057 1.0000 12.500 1.2528 0.03952 0.03290 -0.0090 0.0059 1.0000 13.000 1.2409 0.04627 0.03985 -0.0086 0.0061 1.0000 13.500 1.2372 0.05268 0.04644 -0.0091 0.0064 1.0000 14.000 1.2376 0.05896 0.05293 -0.0099 0.0068 1.0000 14.500 1.2274 0.06678 0.06095 -0.0114 0.0073 1.0000 15.000 1.2163 0.07448 0.06876 -0.0124 0.0078 1.0000 15.500 1.2193 0.08064 0.07505 -0.0133 0.0083 1.0000 16.000 1.2246 0.08628 0.08091 -0.0136 0.0093 1.0000 16.500 1.2386 0.09027 0.08502 -0.0126 0.0105 1.0000