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NACA 1410 (naca1410-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: NACA 1410 (naca1410-il)
Reynolds number: 50,000
Max Cl/Cd: 31.19 at α=5.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-naca1410-il-50000.txt
Download as CSV file: xf-naca1410-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NACA 1410                                       
  
 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
  ------ -------- --------- --------- -------- -------- --------
  -8.250  -0.5190   0.09606   0.08875   0.0124   1.0000   0.3712
  -8.000  -0.5138   0.09287   0.08559   0.0130   1.0000   0.3877
  -7.750  -0.6636   0.06363   0.05616  -0.0284   1.0000   0.1667
  -7.500  -0.6549   0.05879   0.05114  -0.0284   1.0000   0.1638
  -7.250  -0.6487   0.05377   0.04587  -0.0285   1.0000   0.1604
  -7.000  -0.6436   0.04872   0.04037  -0.0283   1.0000   0.1576
  -6.750  -0.6330   0.04499   0.03623  -0.0274   1.0000   0.1602
  -6.500  -0.6208   0.04144   0.03215  -0.0264   1.0000   0.1632
  -6.250  -0.6054   0.03812   0.02822  -0.0252   1.0000   0.1650
  -6.000  -0.5863   0.03523   0.02503  -0.0241   1.0000   0.1682
  -5.750  -0.5653   0.03316   0.02282  -0.0230   1.0000   0.1753
  -5.500  -0.5446   0.03113   0.02049  -0.0219   1.0000   0.1851
  -5.250  -0.5221   0.02923   0.01822  -0.0208   1.0000   0.1956
  -5.000  -0.4987   0.02742   0.01639  -0.0197   1.0000   0.2090
  -4.750  -0.4761   0.02588   0.01491  -0.0186   1.0000   0.2303
  -4.500  -0.4539   0.02443   0.01356  -0.0173   1.0000   0.2622
  -4.250  -0.4321   0.02285   0.01222  -0.0160   1.0000   0.3107
  -4.000  -0.4116   0.02133   0.01122  -0.0143   1.0000   0.3819
  -3.750  -0.3947   0.02040   0.01086  -0.0115   1.0000   0.4712
  -3.500  -0.3790   0.01945   0.01046  -0.0079   1.0000   0.5738
  -3.250  -0.3655   0.01883   0.01038  -0.0029   1.0000   0.6774
  -3.000  -0.3522   0.01862   0.01049   0.0030   1.0000   0.7764
  -2.750  -0.3183   0.01900   0.01088   0.0063   1.0000   0.8739
  -2.500  -0.1443   0.01990   0.01078  -0.0174   1.0000   0.9741
  -2.250  -0.0704   0.01907   0.00959  -0.0285   1.0000   1.0000
  -2.000  -0.0733   0.01848   0.00897  -0.0255   1.0000   1.0000
  -1.750  -0.0774   0.01804   0.00848  -0.0219   1.0000   1.0000
  -1.500  -0.0796   0.01776   0.00813  -0.0181   1.0000   1.0000
  -1.250  -0.0780   0.01762   0.00789  -0.0147   1.0000   1.0000
  -1.000  -0.0726   0.01759   0.00775  -0.0118   1.0000   1.0000
  -0.750  -0.0637   0.01765   0.00769  -0.0094   1.0000   1.0000
  -0.500  -0.0521   0.01779   0.00771  -0.0075   1.0000   1.0000
  -0.250  -0.0384   0.01799   0.00781  -0.0059   1.0000   1.0000
   0.000  -0.0235   0.01825   0.00799  -0.0047   1.0000   1.0000
   0.250  -0.0078   0.01858   0.00824  -0.0037   1.0000   1.0000
   0.500   0.0083   0.01897   0.00858  -0.0029   1.0000   1.0000
   0.750   0.0247   0.01944   0.00900  -0.0023   1.0000   1.0000
   1.000   0.0411   0.01998   0.00951  -0.0019   1.0000   1.0000
   1.250   0.0575   0.02061   0.01011  -0.0017   1.0000   1.0000
   1.500   0.0737   0.02131   0.01081  -0.0016   1.0000   1.0000
   1.750   0.0959   0.02215   0.01166  -0.0029   0.9974   1.0000
   2.000   0.1652   0.02338   0.01298  -0.0125   0.9751   1.0000
   2.250   0.2277   0.02431   0.01403  -0.0204   0.9505   1.0000
   2.500   0.2862   0.02502   0.01490  -0.0270   0.9248   1.0000
   2.750   0.3508   0.02545   0.01553  -0.0340   0.8998   1.0000
   3.000   0.4088   0.02562   0.01594  -0.0391   0.8729   1.0000
   3.250   0.4605   0.02555   0.01611  -0.0424   0.8451   1.0000
   3.500   0.5079   0.02526   0.01604  -0.0442   0.8164   1.0000
   3.750   0.5495   0.02482   0.01580  -0.0442   0.7864   1.0000
   4.000   0.5851   0.02434   0.01549  -0.0429   0.7547   1.0000
   4.250   0.6168   0.02380   0.01506  -0.0406   0.7209   1.0000
   4.500   0.6424   0.02346   0.01478  -0.0374   0.6826   1.0000
   4.750   0.6673   0.02306   0.01435  -0.0339   0.6411   1.0000
   5.000   0.6883   0.02298   0.01420  -0.0303   0.5934   1.0000
   5.250   0.7075   0.02303   0.01410  -0.0266   0.5382   1.0000
   5.500   0.7243   0.02322   0.01397  -0.0227   0.4725   1.0000
   5.750   0.7375   0.02385   0.01417  -0.0189   0.3939   1.0000
   6.000   0.7505   0.02512   0.01489  -0.0156   0.3142   1.0000
   6.250   0.7679   0.02693   0.01620  -0.0134   0.2552   1.0000
   6.500   0.7889   0.02874   0.01772  -0.0120   0.2181   1.0000
   6.750   0.8132   0.03077   0.01975  -0.0110   0.1956   1.0000
   7.000   0.8385   0.03299   0.02203  -0.0102   0.1817   1.0000
   7.250   0.8626   0.03522   0.02428  -0.0095   0.1705   1.0000
   7.500   0.8838   0.03736   0.02669  -0.0084   0.1604   1.0000
   7.750   0.9044   0.04029   0.02995  -0.0074   0.1551   1.0000
   8.000   0.9199   0.04354   0.03377  -0.0059   0.1517   1.0000
   8.250   0.9349   0.04661   0.03720  -0.0045   0.1471   1.0000
   8.500   0.9514   0.04983   0.04050  -0.0036   0.1419   1.0000
   8.750   0.9555   0.05398   0.04523  -0.0018   0.1407   1.0000
   9.000   0.9562   0.05864   0.05037  -0.0003   0.1408   1.0000
   9.250   0.9525   0.06359   0.05571   0.0010   0.1412   1.0000
   9.500   0.9447   0.06874   0.06118   0.0019   0.1417   1.0000
   9.750   0.9350   0.07407   0.06672   0.0024   0.1422   1.0000
  10.000   0.9255   0.07954   0.07237   0.0025   0.1430   1.0000
  10.250   0.8529   0.09051   0.08353  -0.0025   0.1580   1.0000
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