XFOIL Version 6.94 Calculated polar for: GOA 1 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 -0.1837 0.04595 0.01936 -0.0217 1.0000 1.0000 -2.750 -0.1582 0.04571 0.01781 -0.0212 1.0000 1.0000 -2.500 -0.1333 0.04551 0.01650 -0.0207 1.0000 1.0000 -2.250 -0.1086 0.04534 0.01537 -0.0200 1.0000 1.0000 -2.000 -0.0841 0.04521 0.01428 -0.0194 1.0000 1.0000 -1.750 -0.0598 0.04512 0.01347 -0.0187 1.0000 1.0000 -1.500 -0.0357 0.04505 0.01279 -0.0181 1.0000 1.0000 -1.250 -0.0117 0.04502 0.01226 -0.0174 1.0000 1.0000 -1.000 0.0122 0.04501 0.01185 -0.0167 1.0000 1.0000 -0.750 0.0360 0.04504 0.01158 -0.0160 1.0000 1.0000 -0.500 0.0597 0.04509 0.01137 -0.0153 1.0000 1.0000 -0.250 0.0834 0.04517 0.01129 -0.0146 1.0000 1.0000 0.000 0.1072 0.04526 0.01113 -0.0138 1.0000 1.0000 0.250 0.1309 0.04539 0.01127 -0.0131 1.0000 1.0000 0.500 0.1545 0.04555 0.01154 -0.0123 1.0000 1.0000 0.750 0.1782 0.04573 0.01194 -0.0116 1.0000 1.0000 1.000 0.2019 0.04595 0.01249 -0.0109 1.0000 1.0000 1.250 0.2257 0.04618 0.01319 -0.0102 1.0000 1.0000 1.500 0.2499 0.04645 0.01407 -0.0095 1.0000 1.0000 1.750 0.2745 0.04674 0.01514 -0.0089 1.0000 1.0000 2.000 0.3000 0.04706 0.01647 -0.0085 1.0000 1.0000 2.250 0.3269 0.04744 0.01813 -0.0087 1.0000 1.0000 2.500 0.3547 0.04801 0.02029 -0.0099 1.0000 1.0000 2.750 0.3765 0.04924 0.02316 -0.0128 1.0000 1.0000 3.000 0.3822 0.05165 0.02615 -0.0162 1.0000 1.0000 3.250 0.3810 0.05476 0.02908 -0.0191 1.0000 1.0000 3.500 0.3818 0.05787 0.03192 -0.0220 1.0000 1.0000 3.750 0.3852 0.06088 0.03471 -0.0246 1.0000 1.0000 4.000 0.3903 0.06382 0.03749 -0.0271 1.0000 1.0000 4.250 0.3964 0.06673 0.04026 -0.0294 1.0000 1.0000 4.500 0.4033 0.06963 0.04304 -0.0316 1.0000 1.0000 4.750 0.4107 0.07252 0.04584 -0.0336 1.0000 1.0000 5.000 0.4186 0.07543 0.04868 -0.0356 1.0000 1.0000