XFOIL Version 6.94 Calculated polar for: WBL FX 05-H-126 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.045 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 0.0495 0.04233 0.03301 -0.0275 0.3247 0.1664 -2.750 0.0840 0.04045 0.03062 -0.0284 0.3195 0.1595 -2.500 0.1170 0.03908 0.02873 -0.0286 0.3149 0.1573 -2.250 0.1483 0.03798 0.02713 -0.0283 0.3112 0.1566 -2.000 0.1784 0.03711 0.02576 -0.0275 0.3082 0.1556 -1.750 0.2078 0.03627 0.02456 -0.0265 0.3059 0.1564 -1.500 0.2376 0.03545 0.02356 -0.0254 0.3041 0.1594 -1.250 0.2679 0.03491 0.02280 -0.0244 0.3025 0.1673 -1.000 0.2979 0.03436 0.02214 -0.0234 0.3012 0.1793 -0.750 0.3277 0.03393 0.02163 -0.0224 0.3002 0.1958 -0.500 0.3574 0.03336 0.02118 -0.0216 0.2994 0.2331 -0.250 0.3860 0.03032 0.02027 -0.0190 0.2986 1.0001 0.000 0.4161 0.03112 0.02049 -0.0185 0.2974 1.0001 0.250 0.4457 0.03200 0.02102 -0.0184 0.2958 1.0001 0.500 0.4752 0.03298 0.02170 -0.0185 0.2939 1.0001 0.750 0.5044 0.03405 0.02252 -0.0186 0.2921 1.0001 1.000 0.5338 0.03522 0.02347 -0.0189 0.2908 1.0001 1.250 0.5637 0.03641 0.02457 -0.0194 0.2909 1.0001 1.500 0.5938 0.03767 0.02582 -0.0202 0.2917 1.0001 1.750 0.6236 0.03907 0.02726 -0.0212 0.2930 1.0001 2.000 0.6531 0.04066 0.02892 -0.0224 0.2947 1.0001 2.250 0.6817 0.04247 0.03082 -0.0237 0.2968 1.0001 2.500 0.7093 0.04454 0.03298 -0.0251 0.2992 1.0001 2.750 0.7358 0.04683 0.03533 -0.0264 0.3018 1.0001 3.000 0.7611 0.04937 0.03789 -0.0276 0.3045 1.0001 3.250 0.7864 0.05208 0.04051 -0.0283 0.3067 1.0001 3.500 0.7504 0.06593 0.05601 -0.0434 0.3319 1.0001 3.750 0.8057 0.06306 0.05252 -0.0376 0.3264 1.0001