XFOIL Version 6.94 Calculated polar for: mta3 300.2.14.15.2.c13 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.005 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.1120 0.05546 0.03744 -0.0092 1.0000 1.0000 -2.750 -0.1137 0.05369 0.03608 -0.0091 1.0000 1.0000 -2.500 -0.1180 0.05194 0.03477 -0.0087 1.0000 1.0000 -2.250 -0.1238 0.05029 0.03346 -0.0083 1.0000 1.0000 -2.000 -0.1198 0.04877 0.03182 -0.0106 1.0000 1.0000 -1.750 -0.0822 0.04778 0.02943 -0.0190 1.0000 1.0000 -1.500 -0.0345 0.04767 0.02722 -0.0245 1.0000 1.0000 -1.250 0.0017 0.04776 0.02564 -0.0254 1.0000 1.0000 -1.000 0.0313 0.04785 0.02451 -0.0247 1.0000 1.0000 -0.750 0.0579 0.04794 0.02366 -0.0236 1.0000 1.0000 -0.500 0.0829 0.04804 0.02303 -0.0225 1.0000 1.0000 -0.250 0.1073 0.04815 0.02261 -0.0214 1.0000 1.0000 0.000 0.1313 0.04829 0.02233 -0.0203 1.0000 1.0000 0.250 0.1552 0.04846 0.02226 -0.0193 1.0000 1.0000 0.500 0.1792 0.04866 0.02235 -0.0184 1.0000 1.0000 0.750 0.2033 0.04889 0.02262 -0.0176 1.0000 1.0000 1.000 0.2278 0.04915 0.02306 -0.0169 1.0000 1.0000 1.250 0.2528 0.04947 0.02375 -0.0165 1.0000 1.0000 1.500 0.2782 0.04987 0.02474 -0.0164 1.0000 1.0000 1.750 0.3029 0.05049 0.02617 -0.0170 1.0000 1.0000 2.000 0.3224 0.05168 0.02826 -0.0187 1.0000 1.0000 2.250 0.3287 0.05404 0.03112 -0.0213 1.0000 1.0000 2.500 0.3253 0.05723 0.03417 -0.0238 1.0000 1.0000 2.750 0.3238 0.06044 0.03706 -0.0262 1.0000 1.0000 3.000 0.3257 0.06348 0.03980 -0.0285 1.0000 1.0000 3.250 0.3298 0.06640 0.04243 -0.0307 1.0000 1.0000 3.500 0.3353 0.06926 0.04500 -0.0326 1.0000 1.0000 3.750 0.3419 0.07207 0.04755 -0.0345 1.0000 1.0000 4.000 0.3492 0.07487 0.05010 -0.0362 1.0000 1.0000 4.250 0.3571 0.07767 0.05266 -0.0379 1.0000 1.0000 4.500 0.3655 0.08047 0.05523 -0.0395 1.0000 1.0000 4.750 0.3743 0.08329 0.05782 -0.0410 1.0000 1.0000 5.000 0.3833 0.08613 0.06044 -0.0425 1.0000 1.0000