XFOIL         Version 6.94
  
 Calculated polar for: RG 12A-1.8/9.0 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
 ------- -------- --------- --------- -------- ------- -------
   0.000   0.2557   0.00612   0.00206  -0.0537  0.8442  0.8730
   0.500   0.3061   0.00622   0.00203  -0.0518  0.7997  0.8919
   1.000   0.3529   0.00636   0.00204  -0.0492  0.7532  0.9121
   1.500   0.3990   0.00651   0.00206  -0.0466  0.7063  0.9296
   2.000   0.4473   0.00669   0.00208  -0.0447  0.6594  0.9443
   2.500   0.5001   0.00691   0.00215  -0.0440  0.6111  0.9594
   3.000   0.5602   0.00723   0.00233  -0.0450  0.5584  0.9741
   3.500   0.6266   0.00762   0.00255  -0.0476  0.4983  0.9885
   4.000   0.6906   0.00810   0.00284  -0.0501  0.4330  1.0000
   4.500   0.7320   0.00870   0.00318  -0.0477  0.3661  1.0000
   5.000   0.7770   0.00965   0.00367  -0.0463  0.2641  1.0000
   5.500   0.8196   0.01101   0.00438  -0.0446  0.1437  1.0000
   6.000   0.8652   0.01211   0.00517  -0.0433  0.0832  1.0000
   6.500   0.9098   0.01330   0.00611  -0.0417  0.0359  1.0000
   7.000   0.9541   0.01449   0.00717  -0.0401  0.0156  1.0000
   7.500   0.9994   0.01548   0.00819  -0.0387  0.0107  1.0000
   8.000   1.0450   0.01639   0.00924  -0.0372  0.0090  1.0000
   8.500   1.0881   0.01749   0.01048  -0.0354  0.0080  1.0000
   9.000   1.1300   0.01861   0.01180  -0.0334  0.0071  1.0000
   9.500   1.1707   0.01969   0.01305  -0.0314  0.0051  1.0000
  10.000   1.2079   0.02099   0.01455  -0.0288  0.0033  1.0000
  10.500   1.2391   0.02260   0.01633  -0.0254  0.0018  1.0000
  11.000   1.2595   0.02450   0.01848  -0.0204  0.0015  1.0000
  11.500   1.2725   0.02673   0.02100  -0.0147  0.0013  1.0000
  12.000   1.2790   0.02949   0.02409  -0.0090  0.0012  1.0000
  12.500   1.2783   0.03298   0.02793  -0.0040  0.0012  1.0000
  13.000   1.2706   0.03751   0.03281  -0.0003  0.0012  1.0000
  13.500   1.2557   0.04358   0.03925   0.0012  0.0012  1.0000
  14.000   1.2337   0.05181   0.04786  -0.0002  0.0011  1.0000
  14.500   1.2059   0.06263   0.05904  -0.0048  0.0011  1.0000
  15.000   1.1715   0.07651   0.07327  -0.0124  0.0012  1.0000
  15.500   1.1306   0.09365   0.09074  -0.0226  0.0012  1.0000
  16.000   1.0847   0.11371   0.11108  -0.0344  0.0013  1.0000
  16.500   1.0364   0.13595   0.13355  -0.0471  0.0014  1.0000
  17.000   0.9821   0.16184   0.15963  -0.0613  0.0015  1.0000