XFOIL         Version 6.94
  
 Calculated polar for: NASA SC(2)-1010 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   0.8482   0.01321   0.00560  -0.1397  0.0953  0.7723
   1.000   0.8975   0.01424   0.00644  -0.1381  0.0608  0.7852
   1.500   0.9467   0.01527   0.00747  -0.1364  0.0485  0.7980
   2.000   0.9944   0.01645   0.00868  -0.1345  0.0388  0.8100
   2.500   1.0428   0.01741   0.00977  -0.1326  0.0357  0.8209
   3.000   1.0898   0.01855   0.01100  -0.1305  0.0327  0.8323
   3.500   1.1311   0.02106   0.01369  -0.1275  0.0289  0.8432
   4.000   1.1794   0.02209   0.01488  -0.1256  0.0271  0.8545
   4.500   1.2255   0.02414   0.01716  -0.1233  0.0251  0.8653
   5.000   1.2709   0.02648   0.01977  -0.1210  0.0236  0.8760
   5.500   1.3136   0.02854   0.02207  -0.1185  0.0222  0.8877
   6.000   1.3520   0.03162   0.02551  -0.1153  0.0215  0.9005
   6.500   1.3833   0.03516   0.02953  -0.1112  0.0208  0.9148
   7.000   1.4025   0.03998   0.03489  -0.1055  0.0202  0.9328
   7.500   1.3949   0.04734   0.04299  -0.0969  0.0196  1.0001
   8.000   1.3445   0.05741   0.05385  -0.0846  0.0193  1.0001
   8.500   1.2826   0.06750   0.06450  -0.0754  0.0191  1.0001
   9.000   1.2280   0.07879   0.07623  -0.0731  0.0191  1.0001