XFOIL Version 6.94 Calculated polar for: manu4/28 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 -0.0143 0.09782 0.07915 -0.0170 0.9999 1.0001 -2.750 -0.0127 0.09581 0.07743 -0.0162 0.9999 1.0001 -2.500 -0.0118 0.09377 0.07570 -0.0153 0.9999 1.0001 -2.250 -0.0117 0.09172 0.07399 -0.0143 0.9999 1.0001 -2.000 -0.0125 0.08966 0.07223 -0.0132 0.9999 1.0001 -1.750 -0.0145 0.08757 0.07051 -0.0120 0.9999 1.0001 -1.500 -0.0178 0.08547 0.06880 -0.0105 0.9999 1.0001 -1.250 -0.0227 0.08332 0.06704 -0.0089 0.9999 1.0001 -1.000 -0.0284 0.08118 0.06528 -0.0072 0.9999 1.0001 -0.750 -0.0326 0.07921 0.06361 -0.0060 0.9999 1.0001 -0.500 -0.0322 0.07759 0.06218 -0.0057 0.9999 1.0001 -0.250 -0.0252 0.07661 0.06128 -0.0067 0.9999 1.0001 0.000 -0.0121 0.07646 0.06114 -0.0091 0.9999 1.0001 0.250 0.0028 0.07734 0.06201 -0.0124 0.9999 1.0001 0.500 0.0099 0.07950 0.06420 -0.0153 0.9999 1.0001 0.750 0.0064 0.08275 0.06734 -0.0174 0.9999 1.0001 1.000 0.0051 0.08616 0.07043 -0.0199 0.9999 1.0001 1.250 0.0118 0.08945 0.07323 -0.0233 0.9999 1.0001 1.500 0.0250 0.09266 0.07586 -0.0275 0.9999 1.0001 1.750 0.0419 0.09591 0.07843 -0.0319 0.9999 1.0001 2.000 0.0618 0.09919 0.08098 -0.0365 0.9999 1.0001 2.250 0.0829 0.10248 0.08350 -0.0410 0.9999 1.0001 2.500 0.1039 0.10574 0.08599 -0.0451 0.9999 1.0001 2.750 0.1243 0.10896 0.08844 -0.0487 0.9999 1.0001 3.000 0.1440 0.11211 0.09081 -0.0519 0.9999 1.0001 3.250 0.1624 0.11516 0.09314 -0.0546 0.9999 1.0001 3.500 0.1799 0.11814 0.09538 -0.0569 0.9999 1.0001 3.750 0.1964 0.12105 0.09761 -0.0588 0.9999 1.0001 4.000 0.2121 0.12390 0.09982 -0.0606 0.9999 1.0001 4.250 0.2273 0.12667 0.10199 -0.0621 0.9999 1.0001 4.500 0.2421 0.12941 0.10415 -0.0634 0.9999 1.0001 4.750 0.2564 0.13212 0.10630 -0.0646 0.9999 1.0001 5.000 0.2704 0.13479 0.10847 -0.0658 0.9999 1.0001