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 = 0.026 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.4294 0.03653 0.02236 0.0283 1.0000 0.3318 -2.750 -0.4235 0.03642 0.02224 0.0323 1.0000 0.3679 -2.500 -0.4284 0.03593 0.02239 0.0388 1.0000 0.4368 -2.250 -0.4095 0.03547 0.02159 0.0405 1.0000 0.4617 -2.000 -0.3755 0.03506 0.02064 0.0387 1.0000 0.4604 -1.750 -0.3404 0.03467 0.01976 0.0367 1.0000 0.4599 -1.500 -0.3049 0.03432 0.01901 0.0345 1.0000 0.4602 -1.250 -0.2695 0.03402 0.01837 0.0323 1.0000 0.4618 -1.000 -0.2340 0.03371 0.01789 0.0302 1.0000 0.4648 -0.750 -0.1985 0.03338 0.01754 0.0280 1.0000 0.4697 -0.500 -0.1642 0.03311 0.01732 0.0261 1.0000 0.4770 -0.250 -0.1300 0.03278 0.01723 0.0242 1.0000 0.4871 0.000 -0.0969 0.03243 0.01727 0.0226 1.0000 0.5022 0.250 -0.0642 0.03196 0.01744 0.0211 1.0000 0.5279 0.500 0.0890 0.03203 0.01952 -0.0033 1.0000 1.0000 0.750 0.1126 0.03237 0.01994 -0.0027 1.0000 1.0000 1.000 0.1346 0.03271 0.02046 -0.0019 1.0000 1.0000 1.250 0.1555 0.03307 0.02113 -0.0010 1.0000 1.0000 1.500 0.1750 0.03354 0.02202 0.0000 1.0000 1.0000 2.000 0.4302 0.03104 0.01801 -0.0248 0.3200 1.0000 2.250 0.4504 0.03157 0.01841 -0.0229 0.2803 1.0000 2.500 0.4723 0.03199 0.01885 -0.0210 0.2483 1.0000 2.750 0.4949 0.03241 0.01946 -0.0191 0.2068 1.0000 3.000 0.5106 0.03368 0.02005 -0.0166 0.1340 1.0000 3.250 0.5285 0.03496 0.02106 -0.0142 0.1280 1.0000 3.500 0.5486 0.03608 0.02223 -0.0119 0.1257 1.0000 3.750 0.5697 0.03726 0.02356 -0.0097 0.1249 1.0000 4.000 0.5928 0.03853 0.02502 -0.0076 0.1253 1.0000 4.250 0.6197 0.03996 0.02674 -0.0058 0.1264 1.0000 4.500 0.6528 0.04157 0.02874 -0.0047 0.1285 1.0000 4.750 0.6979 0.04370 0.03183 -0.0048 0.1338 1.0000 5.000 0.7459 0.04733 0.03634 -0.0059 0.1409 1.0000