XFOIL Version 6.94 Calculated polar for: kut1 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.030 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.4252 0.03493 0.02087 0.0278 1.0000 0.2749 -2.750 -0.4185 0.03495 0.02095 0.0314 1.0000 0.3049 -2.500 -0.4093 0.03485 0.02087 0.0346 1.0000 0.3370 -2.250 -0.4129 0.03446 0.02115 0.0404 1.0000 0.4055 -2.000 -0.3890 0.03407 0.02040 0.0409 1.0000 0.4224 -1.750 -0.3551 0.03374 0.01955 0.0392 1.0000 0.4211 -1.500 -0.3205 0.03343 0.01881 0.0373 1.0000 0.4202 -1.250 -0.2854 0.03315 0.01818 0.0353 1.0000 0.4201 -1.000 -0.2502 0.03289 0.01769 0.0332 1.0000 0.4211 -0.750 -0.2153 0.03265 0.01734 0.0312 1.0000 0.4232 -0.500 -0.1811 0.03245 0.01712 0.0293 1.0000 0.4266 -0.250 -0.1480 0.03229 0.01700 0.0277 1.0000 0.4314 0.000 -0.1145 0.03205 0.01705 0.0260 1.0000 0.4379 0.250 -0.0830 0.03189 0.01718 0.0247 1.0000 0.4472 0.500 -0.0515 0.03163 0.01745 0.0235 1.0000 0.4605 0.750 -0.0206 0.03129 0.01785 0.0224 1.0000 0.4821 1.000 0.1335 0.03139 0.02067 -0.0020 1.0000 1.0000 1.500 0.3884 0.02935 0.01661 -0.0287 0.3157 1.0000 1.750 0.4088 0.02988 0.01691 -0.0270 0.2730 1.0000 2.000 0.4317 0.03017 0.01715 -0.0254 0.2382 1.0000 2.250 0.4558 0.03037 0.01743 -0.0238 0.2028 1.0000 2.500 0.4726 0.03155 0.01795 -0.0215 0.1198 1.0000 2.750 0.4920 0.03268 0.01905 -0.0193 0.1138 1.0000 3.000 0.5130 0.03367 0.02009 -0.0172 0.1116 1.0000 3.250 0.5345 0.03467 0.02121 -0.0152 0.1110 1.0000 3.500 0.5564 0.03577 0.02243 -0.0131 0.1110 1.0000 3.750 0.5801 0.03697 0.02375 -0.0112 0.1118 1.0000 4.000 0.6083 0.03827 0.02524 -0.0097 0.1136 1.0000 4.250 0.6413 0.03913 0.02666 -0.0084 0.1178 1.0000 4.500 0.6831 0.04063 0.02871 -0.0082 0.1246 1.0000 4.750 0.7305 0.04288 0.03195 -0.0085 0.1332 1.0000 5.000 0.7695 0.04684 0.03681 -0.0085 0.1425 1.0000