XFOIL Version 6.94 Calculated polar for: NACA4412c 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 -0.0681 0.04225 0.03091 -0.0495 0.9999 0.3146 -2.750 -0.0411 0.04111 0.02957 -0.0495 0.9999 0.3310 -2.500 -0.0134 0.04014 0.02842 -0.0496 0.9999 0.3517 -2.250 0.0143 0.03930 0.02751 -0.0496 0.9999 0.3785 -2.000 0.0417 0.03853 0.02685 -0.0494 0.9999 0.4139 -1.750 0.0696 0.03774 0.02637 -0.0493 0.9999 0.4656 -1.500 0.0963 0.03645 0.02607 -0.0485 0.9999 0.5724 -1.250 0.1074 0.03475 0.02516 -0.0442 0.9999 1.0001 -1.000 0.1341 0.03563 0.02525 -0.0445 0.9999 1.0001 -0.750 0.1564 0.03658 0.02586 -0.0444 0.9999 1.0001 -0.500 0.1769 0.03769 0.02678 -0.0444 0.9999 1.0001 -0.250 0.1945 0.03917 0.02823 -0.0446 0.9999 1.0001 0.000 0.4267 0.03309 0.02183 -0.0799 0.8474 1.0001 0.250 0.4991 0.03116 0.01911 -0.0828 0.7204 1.0001 0.500 0.5316 0.03168 0.01893 -0.0817 0.6635 1.0001 0.750 0.5607 0.03252 0.01922 -0.0809 0.6273 1.0001 1.000 0.5892 0.03352 0.01986 -0.0806 0.5995 1.0001 1.250 0.6182 0.03457 0.02061 -0.0807 0.5788 1.0001 1.500 0.6481 0.03572 0.02154 -0.0812 0.5639 1.0001 1.750 0.6782 0.03690 0.02253 -0.0817 0.5518 1.0001 2.000 0.7072 0.03833 0.02396 -0.0827 0.5412 1.0001 2.250 0.7368 0.03960 0.02508 -0.0832 0.5317 1.0001 2.500 0.7649 0.04123 0.02676 -0.0842 0.5224 1.0001 2.750 0.7935 0.04269 0.02816 -0.0849 0.5137 1.0001 3.000 0.8209 0.04444 0.02992 -0.0857 0.5068 1.0001 3.250 0.8459 0.04670 0.03235 -0.0869 0.5018 1.0001 3.500 0.8705 0.04903 0.03478 -0.0880 0.4977 1.0001 3.750 0.8951 0.05131 0.03714 -0.0889 0.4938 1.0001 4.000 0.9214 0.05337 0.03914 -0.0893 0.4898 1.0001 4.250 0.9334 0.05717 0.04324 -0.0903 0.4865 1.0001 4.500 0.9407 0.06142 0.04772 -0.0910 0.4837 1.0001 4.750 0.9411 0.06649 0.05299 -0.0915 0.4831 1.0001 5.000 0.9340 0.07254 0.05921 -0.0921 0.4855 1.0001