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