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 = 0.008 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 0.0119 0.07143 0.05341 -0.0298 1.0001 0.9999 -2.750 -0.0008 0.06977 0.05190 -0.0284 1.0001 0.9999 -2.500 -0.0166 0.06796 0.05024 -0.0268 1.0001 0.9999 -2.250 -0.0364 0.06593 0.04839 -0.0248 1.0001 0.9999 -2.000 -0.0611 0.06359 0.04627 -0.0224 1.0001 0.9999 -1.750 -0.0907 0.06090 0.04381 -0.0196 1.0001 0.9999 -1.500 -0.1132 0.05810 0.04076 -0.0192 1.0001 0.9999 -1.250 -0.0945 0.05716 0.03837 -0.0242 1.0001 0.9999 -1.000 -0.0713 0.05746 0.03753 -0.0260 1.0001 0.9999 -0.750 -0.0503 0.05797 0.03720 -0.0267 1.0001 0.9999 -0.500 -0.0304 0.05860 0.03716 -0.0269 1.0001 0.9999 -0.250 -0.0112 0.05931 0.03731 -0.0270 1.0001 0.9999 0.000 0.0076 0.06010 0.03763 -0.0271 1.0001 0.9999 0.250 0.0260 0.06097 0.03808 -0.0271 1.0001 0.9999 0.500 0.0440 0.06193 0.03869 -0.0271 1.0001 0.9999 0.750 0.0615 0.06298 0.03944 -0.0272 1.0001 0.9999 1.000 0.0785 0.06413 0.04034 -0.0273 1.0001 0.9999 1.250 0.0948 0.06540 0.04139 -0.0275 1.0001 0.9999 1.500 0.1104 0.06680 0.04261 -0.0277 1.0001 0.9999 1.750 0.1250 0.06834 0.04402 -0.0280 1.0001 0.9999 2.000 0.1386 0.07006 0.04563 -0.0284 1.0001 0.9999 2.250 0.1510 0.07196 0.04744 -0.0290 1.0001 0.9999 2.500 0.1622 0.07407 0.04950 -0.0296 1.0001 0.9999 2.750 0.1723 0.07641 0.05178 -0.0304 1.0001 0.9999 3.000 0.1814 0.07893 0.05425 -0.0314 1.0001 0.9999 3.250 0.1899 0.08159 0.05686 -0.0325 1.0001 0.9999 3.500 0.1982 0.08434 0.05955 -0.0336 1.0001 0.9999 3.750 0.2066 0.08713 0.06227 -0.0348 1.0001 0.9999 4.000 0.2154 0.08994 0.06500 -0.0360 1.0001 0.9999 4.250 0.2245 0.09275 0.06774 -0.0372 1.0001 0.9999 4.500 0.2340 0.09556 0.07046 -0.0384 1.0001 0.9999 4.750 0.2440 0.09835 0.07318 -0.0397 1.0001 0.9999 5.000 0.2542 0.10114 0.07588 -0.0409 1.0001 0.9999