XFOIL         Version 6.94
  
 Calculated polar for: RG-15 8.9%                                      
  
 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   0.2402   0.00726   0.00286  -0.0549  0.8880  0.9003
   0.500   0.2940   0.00710   0.00267  -0.0532  0.8569  0.9342
   1.000   0.3614   0.00707   0.00252  -0.0548  0.8216  0.9653
   1.500   0.4431   0.00711   0.00244  -0.0600  0.7785  0.9861
   2.000   0.5099   0.00724   0.00237  -0.0626  0.7289  1.0000
   2.500   0.5497   0.00753   0.00245  -0.0595  0.6781  1.0000
   3.000   0.5969   0.00790   0.00261  -0.0577  0.6227  1.0000
   3.500   0.6451   0.00837   0.00285  -0.0561  0.5633  1.0000
   4.000   0.6932   0.00891   0.00321  -0.0546  0.4994  1.0000
   4.500   0.7412   0.00954   0.00363  -0.0531  0.4325  1.0000
   5.000   0.7880   0.01032   0.00418  -0.0515  0.3566  1.0000
   5.500   0.8345   0.01123   0.00483  -0.0500  0.2790  1.0000
   6.500   0.9229   0.01377   0.00669  -0.0467  0.1212  1.0000
   7.000   0.9664   0.01519   0.00795  -0.0450  0.0700  1.0000
   7.500   1.0063   0.01703   0.00967  -0.0426  0.0359  1.0000
   9.000   1.1023   0.02589   0.01929  -0.0315  0.0139  1.0000
   9.500   1.1321   0.03015   0.02400  -0.0280  0.0133  1.0000
  10.000   1.1561   0.03421   0.02849  -0.0242  0.0124  1.0000
  10.500   1.1692   0.03766   0.03220  -0.0198  0.0112  1.0000
  11.000   1.1586   0.04349   0.03852  -0.0135  0.0109  1.0000
  11.500   1.1363   0.04996   0.04549  -0.0088  0.0108  1.0000
  12.000   1.1087   0.05759   0.05357  -0.0075  0.0108  1.0000
  12.500   1.0776   0.06734   0.06373  -0.0106  0.0109  1.0000
  13.000   1.0406   0.08076   0.07757  -0.0187  0.0112  1.0000
  14.000   0.8833   0.14404   0.14142  -0.0568  0.0160  1.0000