XFOIL Version 6.94 Calculated polar for: manu4/38 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 -0.0792 0.07415 0.05211 -0.0202 0.9999 1.0001 -2.750 -0.0819 0.07217 0.05037 -0.0190 0.9999 1.0001 -2.500 -0.0861 0.07016 0.04863 -0.0175 0.9999 1.0001 -2.250 -0.0917 0.06813 0.04686 -0.0159 0.9999 1.0001 -2.000 -0.0976 0.06609 0.04504 -0.0143 0.9999 1.0001 -1.750 -0.1000 0.06419 0.04311 -0.0135 0.9999 1.0001 -1.500 -0.0943 0.06264 0.04132 -0.0142 0.9999 1.0001 -1.250 -0.0754 0.06163 0.03974 -0.0171 0.9999 1.0001 -1.000 -0.0430 0.06123 0.03844 -0.0222 0.9999 1.0001 -0.750 -0.0015 0.06137 0.03744 -0.0282 0.9999 1.0001 -0.500 0.0414 0.06184 0.03664 -0.0335 0.9999 1.0001 -0.250 0.0801 0.06241 0.03613 -0.0370 0.9999 1.0001 0.000 0.1141 0.06299 0.03582 -0.0392 0.9999 1.0001 0.250 0.1446 0.06356 0.03572 -0.0405 0.9999 1.0001 0.500 0.1729 0.06415 0.03578 -0.0413 0.9999 1.0001 0.750 0.1994 0.06479 0.03608 -0.0418 0.9999 1.0001 1.000 0.2244 0.06550 0.03659 -0.0422 0.9999 1.0001 1.250 0.2477 0.06635 0.03737 -0.0426 0.9999 1.0001 1.500 0.2686 0.06741 0.03849 -0.0430 0.9999 1.0001 1.750 0.2853 0.06887 0.04011 -0.0435 0.9999 1.0001 2.000 0.2947 0.07103 0.04249 -0.0441 0.9999 1.0001 2.250 0.2934 0.07421 0.04576 -0.0448 0.9999 1.0001 2.500 0.2878 0.07787 0.04929 -0.0454 0.9999 1.0001 2.750 0.2852 0.08136 0.05254 -0.0462 0.9999 1.0001 3.000 0.2864 0.08457 0.05548 -0.0471 0.9999 1.0001 3.250 0.2901 0.08757 0.05820 -0.0479 0.9999 1.0001 3.500 0.2955 0.09045 0.06081 -0.0488 0.9999 1.0001 3.750 0.3021 0.09324 0.06332 -0.0496 0.9999 1.0001 4.000 0.3097 0.09595 0.06578 -0.0504 0.9999 1.0001 4.250 0.3178 0.09864 0.06821 -0.0512 0.9999 1.0001 4.500 0.3265 0.10127 0.07059 -0.0519 0.9999 1.0001 4.750 0.3357 0.10389 0.07297 -0.0526 0.9999 1.0001 5.000 0.3452 0.10647 0.07532 -0.0534 0.9999 1.0001