XFOIL         Version 6.94
  
 Calculated polar for: NASA SC(2)-0710 AIRFOIL                         
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.080     Re =     0.450 e 6     Ncrit =   9.000
  
  alpha     CL        CD       CDp       CM    Top_Xtr Bot_Xtr
 ------- -------- --------- --------- -------- ------- -------
   0.500   0.5688   0.00835   0.00416  -0.1175  0.8105  0.7411
   1.000   0.5975   0.01098   0.00476  -0.1115  0.3714  0.7449
   1.500   0.6445   0.01275   0.00542  -0.1104  0.1321  0.7505
   2.000   0.6950   0.01371   0.00608  -0.1092  0.0784  0.7539
   2.500   0.7491   0.01447   0.00671  -0.1089  0.0595  0.7591
   3.000   0.8021   0.01506   0.00738  -0.1081  0.0518  0.7629
   4.500   0.9561   0.01803   0.01048  -0.1050  0.0306  0.7758
   6.000   1.1043   0.02377   0.01689  -0.1005  0.0240  0.7901
   6.500   1.1494   0.02684   0.02036  -0.0981  0.0235  0.7935
   7.000   1.1938   0.02924   0.02305  -0.0962  0.0223  0.7987
   7.500   1.2301   0.03322   0.02753  -0.0929  0.0219  0.8026
   8.000   1.2567   0.03832   0.03323  -0.0885  0.0216  0.8071
   8.500   1.2714   0.04412   0.03968  -0.0830  0.0212  0.8114
   9.000   1.2658   0.05114   0.04731  -0.0757  0.0208  0.8154
   9.500   1.2302   0.05908   0.05581  -0.0661  0.0206  0.8204
  10.000   1.1754   0.06801   0.06522  -0.0582  0.0205  0.8244