XFOIL         Version 6.94
  
 Calculated polar for: RAF 15 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   0.0977   0.01017   0.00359  -0.0100  0.9302  0.1887
   1.000   0.5179   0.00777   0.00277  -0.0768  0.7175  1.0000
   1.500   0.5481   0.00869   0.00280  -0.0708  0.5480  1.0000
   2.000   0.5852   0.00933   0.00293  -0.0667  0.4563  1.0000
   2.500   0.6268   0.00982   0.00313  -0.0637  0.4018  1.0000
   3.000   0.6702   0.01026   0.00338  -0.0610  0.3632  1.0000
   3.500   0.7090   0.01093   0.00357  -0.0576  0.2572  1.0000
   4.000   0.7512   0.01158   0.00388  -0.0548  0.1808  1.0000
   4.500   0.7899   0.01271   0.00447  -0.0516  0.0879  1.0000
   5.000   0.8299   0.01378   0.00524  -0.0485  0.0378  1.0000
   5.500   0.8735   0.01446   0.00596  -0.0459  0.0286  1.0000
   6.000   0.9169   0.01517   0.00672  -0.0432  0.0189  1.0000
   6.500   0.9577   0.01618   0.00780  -0.0399  0.0066  1.0000
   7.000   0.9991   0.01712   0.00897  -0.0367  0.0071  1.0000
   7.500   1.0400   0.01813   0.01022  -0.0334  0.0078  1.0000
   8.000   1.0776   0.01949   0.01189  -0.0294  0.0087  1.0000
   8.500   1.1117   0.02113   0.01388  -0.0248  0.0098  1.0000
   9.000   1.1339   0.02367   0.01672  -0.0185  0.0102  1.0000
   9.500   1.1477   0.02696   0.02026  -0.0112  0.0106  1.0000
  10.000   1.1639   0.03056   0.02423  -0.0045  0.0111  1.0000
  10.500   1.1628   0.03669   0.03113   0.0039  0.0123  1.0000
  11.000   1.1514   0.04297   0.03779   0.0120  0.0131  1.0000