XFOIL Version 6.94 Calculated polar for: manu01 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.012 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.4347 0.07357 0.06179 0.0450 0.9999 0.6651 -2.750 -0.4188 0.07125 0.05935 0.0416 0.9999 0.6485 -2.500 -0.3978 0.06909 0.05700 0.0378 0.9999 0.6356 -2.250 -0.3723 0.06684 0.05460 0.0342 0.9999 0.6271 -2.000 -0.3437 0.06508 0.05263 0.0299 0.9999 0.6203 -1.750 -0.3113 0.06301 0.05041 0.0262 0.9999 0.6162 -1.500 -0.2759 0.06124 0.04847 0.0219 0.9999 0.6122 -1.250 -0.2376 0.05970 0.04672 0.0174 0.9999 0.6090 -1.000 -0.1974 0.05825 0.04513 0.0129 0.9999 0.6097 -0.750 -0.1561 0.05691 0.04369 0.0085 0.9999 0.6187 -0.500 -0.1132 0.05531 0.04220 0.0044 0.9999 0.6342 -0.250 -0.0662 0.05390 0.04095 -0.0007 0.9999 0.6547 0.000 -0.0175 0.05255 0.04008 -0.0067 0.9999 0.6840 0.250 0.0220 0.05207 0.04045 -0.0128 0.9999 0.7273 0.500 0.0118 0.05615 0.04495 -0.0152 0.9999 0.7323 0.750 0.2822 0.05964 0.04692 -0.0743 0.8355 1.0001 1.000 0.3434 0.06190 0.04799 -0.0832 0.8009 1.0001 1.250 0.3826 0.06417 0.04946 -0.0877 0.7787 1.0001 1.500 0.4154 0.06651 0.05118 -0.0910 0.7629 1.0001 1.750 0.4404 0.06907 0.05326 -0.0931 0.7513 1.0001 2.000 0.4560 0.07199 0.05577 -0.0941 0.7436 1.0001 2.250 0.4643 0.07516 0.05864 -0.0942 0.7387 1.0001 2.500 0.4753 0.07831 0.06146 -0.0948 0.7352 1.0001 2.750 0.4863 0.08149 0.06435 -0.0953 0.7329 1.0001 3.000 0.4964 0.08471 0.06730 -0.0957 0.7314 1.0001 3.250 0.5032 0.08801 0.07036 -0.0958 0.7309 1.0001 3.500 0.5054 0.09138 0.07353 -0.0953 0.7320 1.0001 3.750 0.5056 0.09474 0.07670 -0.0947 0.7345 1.0001 4.000 0.5072 0.09806 0.07983 -0.0943 0.7375 1.0001 4.250 0.5113 0.10137 0.08296 -0.0941 0.7415 1.0001 4.500 0.5199 0.10474 0.08613 -0.0946 0.7456 1.0001 4.750 0.5114 0.10789 0.08916 -0.0931 0.7553 1.0001 5.000 0.5152 0.11128 0.09239 -0.0932 0.7646 1.0001