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NACA 65(1)-412 (naca651412-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: NACA 65(1)-412 (naca651412-il)
Reynolds number: 50,000
Max Cl/Cd: 28.28 at α=7°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-naca651412-il-50000.txt
Download as CSV file: xf-naca651412-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NACA 65(1)-412                                  
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.050 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -10.500  -0.4057   0.12005   0.11281  -0.0311   1.0000   0.2626
 -10.250  -0.4127   0.11846   0.11130  -0.0312   1.0000   0.2753
 -10.000  -0.4073   0.11537   0.10825  -0.0304   1.0000   0.2902
  -9.750  -0.4068   0.11281   0.10576  -0.0297   1.0000   0.3055
  -9.500  -0.3812   0.10789   0.10081  -0.0281   1.0000   0.3246
  -9.250  -0.3736   0.10490   0.09787  -0.0270   1.0000   0.3415
  -9.000  -0.3681   0.10215   0.09517  -0.0258   1.0000   0.3585
  -8.750  -0.3665   0.09981   0.09289  -0.0244   1.0000   0.3768
  -8.500  -0.3837   0.09909   0.09232  -0.0223   1.0000   0.3955
  -8.000  -0.3493   0.09182   0.08506  -0.0195   1.0000   0.4333
  -7.750  -0.3627   0.09098   0.08437  -0.0163   1.0000   0.4556
  -7.500  -0.3402   0.08708   0.08045  -0.0152   1.0000   0.4756
  -5.750  -0.5880   0.05337   0.04607  -0.0341   1.0000   0.1721
  -5.500  -0.5694   0.04895   0.04118  -0.0343   1.0000   0.1526
  -5.250  -0.5492   0.04519   0.03663  -0.0346   1.0000   0.1392
  -5.000  -0.5289   0.04199   0.03305  -0.0344   1.0000   0.1336
  -4.750  -0.5067   0.03941   0.02995  -0.0342   1.0000   0.1318
  -4.500  -0.4840   0.03724   0.02736  -0.0338   1.0000   0.1320
  -4.250  -0.4602   0.03527   0.02495  -0.0333   1.0000   0.1313
  -4.000  -0.4366   0.03359   0.02292  -0.0325   1.0000   0.1312
  -3.750  -0.4139   0.03217   0.02125  -0.0316   1.0000   0.1336
  -3.500  -0.3917   0.03113   0.01996  -0.0305   1.0000   0.1398
  -3.250  -0.3709   0.02997   0.01884  -0.0290   1.0000   0.1460
  -3.000  -0.3505   0.02915   0.01796  -0.0273   1.0000   0.1528
  -2.750  -0.3308   0.02830   0.01713  -0.0258   1.0000   0.1643
  -2.500  -0.3103   0.02748   0.01633  -0.0249   1.0000   0.1857
  -2.250  -0.1479   0.02880   0.02034  -0.0268   1.0000   1.0000
  -2.000  -0.1499   0.02853   0.01992  -0.0240   1.0000   1.0000
  -1.750  -0.1520   0.02824   0.01950  -0.0211   1.0000   1.0000
  -1.500  -0.1542   0.02791   0.01905  -0.0182   1.0000   1.0000
  -1.250  -0.1567   0.02756   0.01858  -0.0153   1.0000   1.0000
  -1.000  -0.1591   0.02717   0.01808  -0.0123   1.0000   1.0000
  -0.750  -0.1611   0.02676   0.01756  -0.0094   1.0000   1.0000
  -0.500  -0.1615   0.02636   0.01704  -0.0066   1.0000   1.0000
  -0.250  -0.1562   0.02611   0.01665  -0.0049   1.0000   1.0000
   0.000  -0.1434   0.02610   0.01647  -0.0044   1.0000   1.0000
   0.250  -0.1257   0.02628   0.01648  -0.0046   1.0000   1.0000
   0.500  -0.1054   0.02660   0.01663  -0.0053   1.0000   1.0000
   0.750  -0.0838   0.02703   0.01688  -0.0062   1.0000   1.0000
   1.000  -0.0616   0.02753   0.01723  -0.0071   1.0000   1.0000
   1.250  -0.0392   0.02809   0.01766  -0.0080   1.0000   1.0000
   1.500  -0.0168   0.02872   0.01817  -0.0089   1.0000   1.0000
   1.750   0.0053   0.02938   0.01872  -0.0097   1.0000   1.0000
   2.000   0.0270   0.03010   0.01935  -0.0105   1.0000   1.0000
   2.250   0.0485   0.03086   0.02004  -0.0112   1.0000   1.0000
   2.500   0.0696   0.03166   0.02079  -0.0119   1.0000   1.0000
   2.750   0.0904   0.03251   0.02159  -0.0126   1.0000   1.0000
   3.000   0.1107   0.03340   0.02245  -0.0132   1.0000   1.0000
   3.250   0.1308   0.03434   0.02338  -0.0138   1.0000   1.0000
   3.500   0.1504   0.03533   0.02437  -0.0144   1.0000   1.0000
   3.750   0.1696   0.03636   0.02542  -0.0150   1.0000   1.0000
   4.000   0.1886   0.03746   0.02654  -0.0156   1.0000   1.0000
   4.250   0.2070   0.03860   0.02772  -0.0161   1.0000   1.0000
   4.500   0.2250   0.03981   0.02898  -0.0167   1.0000   1.0000
   4.750   0.2426   0.04109   0.03033  -0.0174   1.0000   1.0000
   5.000   0.2597   0.04244   0.03176  -0.0180   1.0000   1.0000
   5.250   0.2764   0.04387   0.03327  -0.0187   1.0000   1.0000
   5.500   0.2925   0.04538   0.03488  -0.0194   1.0000   1.0000
   5.750   0.4974   0.05046   0.04069  -0.0463   0.8339   1.0000
   6.000   0.5298   0.05158   0.04204  -0.0477   0.8108   1.0000
   6.250   0.5659   0.05260   0.04333  -0.0493   0.7858   1.0000
   6.500   0.6097   0.05323   0.04433  -0.0511   0.7583   1.0000
   6.750   0.6595   0.05275   0.04428  -0.0520   0.7233   1.0000
   7.000   0.8225   0.02908   0.01911  -0.0301   0.2324   1.0000
   7.250   0.8204   0.03184   0.02119  -0.0267   0.1833   1.0000
   7.500   0.8413   0.03388   0.02294  -0.0251   0.1551   1.0000
   7.750   0.9190   0.03635   0.02521  -0.0298   0.1275   1.0000
   8.000   0.9822   0.03959   0.02850  -0.0337   0.1160   1.0000
   8.250   1.0202   0.04252   0.03166  -0.0348   0.1099   1.0000
   8.500   1.0545   0.04612   0.03551  -0.0355   0.1080   1.0000
   8.750   1.0804   0.04988   0.03962  -0.0352   0.1073   1.0000
   9.000   1.0992   0.05395   0.04402  -0.0343   0.1063   1.0000
   9.250   1.1087   0.05728   0.04786  -0.0322   0.1058   1.0000
   9.500   1.1190   0.06129   0.05228  -0.0306   0.1061   1.0000
   9.750   1.1108   0.06417   0.05591  -0.0267   0.1087   1.0000
  10.000   1.0938   0.06828   0.06060  -0.0230   0.1124   1.0000
  10.250   1.0815   0.07268   0.06539  -0.0206   0.1155   1.0000
  10.500   1.0815   0.07781   0.07069  -0.0195   0.1187   1.0000
  10.750   1.0582   0.08150   0.07467  -0.0167   0.1221   1.0000
  11.000   1.0112   0.08572   0.07911  -0.0148   0.1239   1.0000
  11.250   0.9667   0.09185   0.08541  -0.0160   0.1261   1.0000
  11.500   0.9325   0.09953   0.09315  -0.0192   0.1300   1.0000
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