XFOIL Version 6.94 Calculated polar for: nabab 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.018 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.3722 0.05368 0.04127 -0.0086 0.9999 0.2122 -2.750 -0.3437 0.05092 0.03803 -0.0089 0.9999 0.2062 -2.500 -0.3121 0.04849 0.03497 -0.0092 0.9999 0.2014 -2.250 -0.2789 0.04639 0.03222 -0.0094 0.9999 0.1988 -2.000 -0.2441 0.04420 0.02972 -0.0097 0.9999 0.1992 -1.750 -0.2080 0.04207 0.02748 -0.0103 0.9999 0.2039 -1.500 -0.1711 0.04030 0.02560 -0.0108 0.9999 0.2137 -1.250 -0.1350 0.03855 0.02400 -0.0112 0.9999 0.2297 -1.000 -0.1004 0.03686 0.02264 -0.0112 0.9999 0.2568 -0.750 -0.0664 0.03466 0.02147 -0.0113 0.9999 0.3155 -0.500 0.0170 0.03151 0.01981 -0.0182 0.9999 1.0001 -0.250 0.0451 0.03153 0.01960 -0.0170 0.9999 1.0001 0.000 0.0746 0.03152 0.01961 -0.0162 0.9999 1.0001 0.250 0.1053 0.03153 0.01979 -0.0161 0.9999 1.0001 0.500 0.1352 0.03168 0.02021 -0.0164 0.9999 1.0001 0.750 0.2449 0.03101 0.01954 -0.0316 0.7696 1.0001 1.000 0.3019 0.03127 0.01869 -0.0343 0.6534 1.0001 1.250 0.3332 0.03199 0.01885 -0.0340 0.6332 1.0001 1.500 0.3617 0.03279 0.01926 -0.0335 0.6174 1.0001 1.750 0.3805 0.03417 0.01987 -0.0306 0.5700 1.0001 2.000 0.3943 0.03569 0.02064 -0.0270 0.5125 1.0001 2.250 0.4111 0.03675 0.02122 -0.0244 0.4707 1.0001 2.500 0.4307 0.03751 0.02161 -0.0224 0.4416 1.0001 2.750 0.4517 0.03809 0.02193 -0.0207 0.4181 1.0001 3.000 0.4734 0.03862 0.02219 -0.0191 0.3975 1.0001 3.250 0.4948 0.03917 0.02258 -0.0176 0.3754 1.0001 3.500 0.5164 0.03979 0.02307 -0.0160 0.3523 1.0001 3.750 0.5403 0.04054 0.02377 -0.0148 0.3344 1.0001 4.000 0.5634 0.04147 0.02468 -0.0134 0.3085 1.0001 4.250 0.5872 0.04273 0.02592 -0.0120 0.2809 1.0001 4.500 0.6129 0.04456 0.02771 -0.0109 0.2543 1.0001 4.750 0.6401 0.04693 0.03036 -0.0102 0.2223 1.0001 5.000 0.6677 0.04988 0.03352 -0.0097 0.1987 1.0001