XFOIL         Version 6.94
  
 Calculated polar for: 20-32C 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   0.7053   0.01204   0.00416  -0.1296  0.5262  0.0290
   0.500   0.7603   0.01225   0.00402  -0.1293  0.4919  0.0363
   1.000   0.8161   0.01221   0.00355  -0.1292  0.4601  0.0360
   1.500   0.8714   0.01230   0.00348  -0.1291  0.4317  0.0623
   2.000   0.9205   0.01102   0.00389  -0.1283  0.4070  1.0000
   2.500   0.9744   0.01155   0.00415  -0.1280  0.3859  1.0000
   3.000   1.0280   0.01211   0.00447  -0.1277  0.3682  1.0000
   3.500   1.0815   0.01265   0.00486  -0.1274  0.3532  1.0000
   4.000   1.1310   0.01324   0.00486  -0.1268  0.2737  1.0000
   4.500   1.1832   0.01382   0.00529  -0.1264  0.2524  1.0000
   5.000   1.2163   0.01709   0.00730  -0.1240  0.0524  1.0000
   5.500   1.2640   0.01820   0.00830  -0.1228  0.0080  1.0000
   6.000   1.3134   0.01896   0.00928  -0.1218  0.0085  1.0000
   6.500   1.3617   0.01984   0.01041  -0.1204  0.0093  1.0000
   7.000   1.4067   0.02107   0.01199  -0.1185  0.0102  1.0000
   7.500   1.4480   0.02257   0.01382  -0.1161  0.0122  1.0000
   8.000   1.4820   0.02468   0.01635  -0.1125  0.0152  1.0000
   8.500   1.4894   0.02835   0.02042  -0.1060  0.0171  1.0000
   9.000   1.5069   0.03069   0.02304  -0.1006  0.0217  1.0000
   9.500   1.4889   0.03649   0.02911  -0.0945  0.0241  1.0000