XFOIL Version 6.94 Calculated polar for: lb3000 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.022 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.2947 0.05694 0.04529 -0.0234 1.0001 0.3835 -2.750 -0.2664 0.05503 0.04303 -0.0261 1.0001 0.3853 -2.500 -0.2368 0.05347 0.04103 -0.0288 1.0001 0.3907 -2.250 -0.2124 0.05234 0.03984 -0.0293 1.0001 0.4004 -2.000 -0.1834 0.05137 0.03856 -0.0310 1.0001 0.4128 -1.750 -0.1568 0.05053 0.03765 -0.0317 1.0001 0.4252 -1.500 -0.1277 0.04982 0.03674 -0.0330 1.0001 0.4406 -1.250 -0.1015 0.04926 0.03621 -0.0334 1.0001 0.4644 -1.000 -0.0762 0.04873 0.03588 -0.0334 1.0001 0.5003 -0.750 -0.0513 0.04806 0.03563 -0.0330 1.0001 0.5552 -0.500 -0.0340 0.04651 0.03536 -0.0292 1.0001 0.6906 -0.250 -0.0048 0.04507 0.03317 -0.0310 1.0001 0.9999 0.000 0.0179 0.04604 0.03351 -0.0314 1.0001 0.9999 0.250 0.0383 0.04710 0.03415 -0.0316 1.0001 0.9999 0.500 0.0578 0.04826 0.03503 -0.0319 1.0001 0.9999 0.750 0.0762 0.04956 0.03611 -0.0322 1.0001 0.9999 1.000 0.0933 0.05104 0.03743 -0.0327 1.0001 0.9999 1.250 0.1087 0.05276 0.03905 -0.0333 1.0001 0.9999 1.500 0.1214 0.05484 0.04108 -0.0341 1.0001 0.9999 1.750 0.1302 0.05744 0.04369 -0.0352 1.0001 0.9999 2.000 0.1873 0.06118 0.04734 -0.0464 0.9713 0.9999 2.250 0.3112 0.06474 0.05065 -0.0665 0.8887 0.9999 2.500 0.3666 0.06724 0.05303 -0.0730 0.8336 0.9999 2.750 0.4110 0.06966 0.05535 -0.0771 0.7931 0.9999 3.000 0.4572 0.07166 0.05722 -0.0805 0.7588 0.9999 3.250 0.4687 0.07438 0.05988 -0.0801 0.7352 0.9999 3.500 0.4861 0.07719 0.06262 -0.0807 0.7175 0.9999 3.750 0.5073 0.07981 0.06516 -0.0814 0.7009 0.9999 4.000 0.5178 0.08267 0.06797 -0.0812 0.6867 0.9999 4.250 0.5173 0.08600 0.07124 -0.0803 0.6765 0.9999 4.500 0.5291 0.08924 0.07443 -0.0808 0.6689 0.9999 4.750 0.5439 0.09269 0.07782 -0.0818 0.6640 0.9999 5.000 0.5404 0.09655 0.08166 -0.0814 0.6628 0.9999