XFOIL         Version 6.94
  
 Calculated polar for: pp12                                            
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.045 e 6     Ncrit =   9.000
  
  alpha     CL        CD       CDp       CM    Top_Xtr Bot_Xtr
 ------- -------- --------- --------- -------- ------- -------
   0.000  -0.0503   0.04460   0.03862  -0.0013  0.9999  0.6740
   0.250  -0.0320   0.04255   0.03665   0.0003  0.9999  0.7184
   0.500  -0.0033   0.04124   0.03537  -0.0010  0.9999  0.7510
   0.750   0.0378   0.04066   0.03479  -0.0055  0.9999  0.7763
   1.000   0.0937   0.04070   0.03480  -0.0141  0.9999  0.7825
   1.250   0.1514   0.04196   0.03601  -0.0242  0.9999  0.7528
   1.500   0.2077   0.04525   0.03901  -0.0368  0.9999  0.6086
   1.750   0.7069   0.03409   0.02193  -0.0988  0.1777  0.2190
   2.000   0.7412   0.03509   0.02260  -0.1004  0.1687  0.2167
   2.250   0.7881   0.03586   0.02317  -0.1036  0.1614  0.2169
   2.500   0.8431   0.03682   0.02380  -0.1084  0.1520  0.2231
   2.750   0.9911   0.03776   0.02475  -0.1297  0.1458  0.2769
   3.000   1.1137   0.03929   0.02686  -0.1461  0.1454  1.0001
   3.250   1.1551   0.04147   0.02885  -0.1474  0.1467  1.0001
   3.500   1.1918   0.04405   0.03142  -0.1484  0.1484  1.0001
   3.750   1.2203   0.04522   0.03292  -0.1468  0.1515  1.0001
   4.000   1.2463   0.04714   0.03526  -0.1453  0.1548  1.0001
   4.250   1.2708   0.04976   0.03810  -0.1440  0.1538  1.0001
   4.500   1.2932   0.05243   0.04126  -0.1420  0.1600  1.0001
   4.750   1.3163   0.05643   0.04551  -0.1407  0.1661  1.0001
   5.000   1.3357   0.05838   0.04797  -0.1383  0.1738  1.0001