XFOIL         Version 6.94
  
 Calculated polar for: MH 122  9.32%                                   
  
 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   0.4602   0.01173   0.00652  -0.1273  0.8975  0.7308
   1.000   0.5846   0.01114   0.00631  -0.1295  0.8857  0.8222
   1.500   0.6356   0.01064   0.00611  -0.1276  0.8763  0.9091
   2.000   0.7117   0.00997   0.00550  -0.1312  0.8653  1.0000
   2.500   0.7776   0.00932   0.00487  -0.1322  0.8459  1.0000
   3.000   0.8361   0.00900   0.00460  -0.1320  0.8253  1.0000
   3.500   0.8905   0.00877   0.00447  -0.1310  0.7996  1.0000
   4.000   0.9382   0.00863   0.00441  -0.1285  0.7572  1.0000
   4.500   0.9776   0.00891   0.00435  -0.1241  0.6311  1.0000
   5.000   0.9817   0.01146   0.00546  -0.1139  0.3695  1.0000
   5.500   0.9987   0.01401   0.00693  -0.1075  0.1669  1.0000
   6.000   1.0293   0.01595   0.00840  -0.1033  0.0837  1.0000
   6.500   1.0614   0.01776   0.01011  -0.0993  0.0505  1.0000
   7.000   1.0971   0.01938   0.01178  -0.0959  0.0276  1.0000
   7.500   1.1255   0.02210   0.01452  -0.0914  0.0095  1.0000
   8.000   1.1578   0.02503   0.01774  -0.0874  0.0069  1.0000
   8.500   1.1942   0.02888   0.02194  -0.0845  0.0060  1.0000
   9.000   1.2322   0.03488   0.02866  -0.0819  0.0058  1.0000
   9.500   1.2437   0.04370   0.03862  -0.0751  0.0064  1.0000
  10.000   1.2096   0.05437   0.05030  -0.0640  0.0075  1.0000
  10.500   1.1722   0.06274   0.05921  -0.0565  0.0080  1.0000
  11.000   1.1302   0.07219   0.06909  -0.0528  0.0083  1.0000
  11.500   1.0875   0.08347   0.08072  -0.0539  0.0084  1.0000
  12.000   1.0465   0.09793   0.09545  -0.0611  0.0084  1.0000
  12.500   1.0106   0.12124   0.11898  -0.0771  0.0085  1.0000