XFOIL Version 6.96
Calculated polar for: USA 5 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.025 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-1.500 -0.1435 0.04707 0.03653 -0.0147 1.0000 0.2129
-1.000 -0.1201 0.04327 0.03267 -0.0091 1.0000 0.2953
-0.500 -0.0642 0.03743 0.02573 -0.0111 1.0000 0.3143
0.000 -0.0046 0.03431 0.02190 -0.0134 1.0000 0.3686
0.500 0.0569 0.03276 0.02001 -0.0173 1.0000 0.4510
1.000 0.1522 0.03264 0.02068 -0.0288 1.0000 1.0000
1.500 0.1811 0.03523 0.02228 -0.0288 1.0000 1.0000
2.000 0.2007 0.03816 0.02460 -0.0287 1.0000 1.0000
2.500 0.2567 0.04177 0.02764 -0.0363 0.9742 1.0000
3.000 0.3312 0.04626 0.03171 -0.0466 0.9319 1.0000
3.500 0.4038 0.05084 0.03602 -0.0555 0.8957 1.0000
4.000 0.4191 0.05296 0.03808 -0.0541 0.8678 1.0000
4.500 0.4568 0.05710 0.04243 -0.0574 0.8378 1.0000
5.000 0.4959 0.06225 0.04756 -0.0614 0.8222 1.0000
5.500 0.4828 0.06583 0.05114 -0.0579 0.8365 1.0000
7.000 0.6635 0.08030 0.06554 -0.0679 0.6878 1.0000
7.500 0.6539 0.08395 0.06925 -0.0653 0.6735 1.0000
8.000 0.6542 0.08908 0.07445 -0.0645 0.6589 1.0000
8.500 0.6543 0.09601 0.08145 -0.0653 0.6627 1.0000
9.000 0.5729 0.10123 0.08681 -0.0627 0.8359 1.0000
9.500 0.6073 0.10984 0.09549 -0.0672 0.8215 1.0000
10.000 0.6555 0.12451 0.11028 -0.0740 0.8085 1.0000