XFOIL         Version 6.94
  
 Calculated polar for: NASA SC(2)-0712 AIRFOIL                         
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.080     Re =     0.300 e 6     Ncrit =   9.000
  
  alpha     CL        CD       CDp       CM    Top_Xtr Bot_Xtr
 ------- -------- --------- --------- -------- ------- -------
   0.000   0.4825   0.01259   0.00813  -0.1105  0.9103  0.7259
   0.500   0.5470   0.01198   0.00754  -0.1113  0.8813  0.7317
   1.000   0.6035   0.01164   0.00712  -0.1098  0.8143  0.7361
   1.500   0.6466   0.01286   0.00697  -0.1063  0.5230  0.7414
   2.000   0.6793   0.01520   0.00785  -0.1023  0.2309  0.7449
   2.500   0.7304   0.01650   0.00851  -0.1020  0.1293  0.7509
   3.000   0.7789   0.01744   0.00928  -0.1004  0.0981  0.7543
   3.500   0.8341   0.01837   0.01009  -0.1007  0.0783  0.7603
   4.000   0.8816   0.01941   0.01112  -0.0988  0.0682  0.7633
   4.500   0.9340   0.02036   0.01209  -0.0983  0.0615  0.7686
   5.500   1.0319   0.02338   0.01521  -0.0957  0.0497  0.7769
   6.000   1.0838   0.02455   0.01657  -0.0949  0.0451  0.7817
   6.500   1.1313   0.02601   0.01812  -0.0933  0.0418  0.7864
   7.000   1.1805   0.03010   0.02244  -0.0925  0.0395  0.7916
   7.500   1.2253   0.03220   0.02489  -0.0904  0.0387  0.7960
   8.000   1.2677   0.03500   0.02811  -0.0883  0.0374  0.8014
   8.500   1.3008   0.03840   0.03209  -0.0847  0.0357  0.8063
   9.000   1.3285   0.04197   0.03610  -0.0809  0.0338  0.8116
   9.500   1.3460   0.04565   0.04019  -0.0758  0.0324  0.8164
  10.000   1.3506   0.05010   0.04504  -0.0697  0.0314  0.8215
  10.500   1.3238   0.05594   0.05140  -0.0601  0.0310  0.8264
  11.000   1.2779   0.06447   0.06051  -0.0526  0.0310  0.8312