XFOIL         Version 6.94
  
 Calculated polar for: USA 29 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   0.4503   0.01027   0.00319  -0.0915  0.5968  0.3161
   0.500   0.4994   0.01060   0.00339  -0.0900  0.5835  0.3362
   1.000   0.5471   0.01078   0.00358  -0.0882  0.5731  0.3515
   1.500   0.5946   0.01104   0.00372  -0.0865  0.5611  0.3648
   2.000   0.6415   0.01121   0.00390  -0.0846  0.5506  0.3770
   2.500   0.6881   0.01130   0.00402  -0.0828  0.5398  0.3875
   3.500   0.7468   0.01138   0.00392  -0.0716  0.4848  0.4033
   4.000   0.7691   0.01164   0.00402  -0.0647  0.4496  0.4101
   4.500   0.7970   0.01197   0.00425  -0.0592  0.4164  0.4160
   5.000   0.8215   0.01249   0.00463  -0.0532  0.3761  0.4231
   5.500   0.8377   0.01343   0.00525  -0.0460  0.3130  0.4290
   6.000   0.8091   0.01628   0.00717  -0.0320  0.1401  0.4332
   6.500   0.7998   0.01865   0.00911  -0.0221  0.0061  0.4400
   7.000   0.8300   0.01949   0.01003  -0.0184  0.0053  0.4479
   7.500   0.8590   0.02043   0.01112  -0.0149  0.0052  0.4574
   8.000   0.8864   0.02156   0.01241  -0.0114  0.0054  0.4682
   9.000   1.1112   0.02656   0.01898  -0.0439  0.0068  1.0000
   9.500   1.1174   0.02956   0.02215  -0.0389  0.0073  1.0000
  10.000   1.1145   0.03362   0.02639  -0.0341  0.0078  1.0000
  11.000   1.1118   0.04286   0.03597  -0.0273  0.0090  1.0000
  11.500   1.0987   0.04907   0.04234  -0.0242  0.0101  1.0000
  12.000   1.1019   0.05337   0.04677  -0.0204  0.0122  1.0000