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NACA 63-210 AIRFOIL (n63210-il) Xfoil prediction polar at RE=50,000 Ncrit=5


Details Polar file
Airfoil: NACA 63-210 AIRFOIL (n63210-il)
Reynolds number: 50,000
Max Cl/Cd: 32 at α=4.75°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-n63210-il-50000-n5.txt
Download as CSV file: xf-n63210-il-50000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NACA 63-210 AIRFOIL                             
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.050 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -10.000  -0.5648   0.09670   0.08964  -0.0205   1.0000   0.0549
  -9.750  -0.5695   0.09094   0.08395  -0.0241   1.0000   0.0546
  -9.500  -0.5765   0.08452   0.07758  -0.0288   1.0000   0.0539
  -9.250  -0.5893   0.07823   0.07133  -0.0336   1.0000   0.0529
  -9.000  -0.6067   0.07286   0.06596  -0.0371   1.0000   0.0522
  -8.750  -0.6218   0.06813   0.06114  -0.0390   1.0000   0.0517
  -8.500  -0.6300   0.06383   0.05669  -0.0402   1.0000   0.0522
  -8.250  -0.6341   0.05976   0.05239  -0.0407   1.0000   0.0532
  -8.000  -0.6342   0.05575   0.04808  -0.0409   1.0000   0.0541
  -7.750  -0.6305   0.05177   0.04373  -0.0407   1.0000   0.0548
  -7.500  -0.6229   0.04803   0.03957  -0.0402   1.0000   0.0554
  -7.250  -0.6124   0.04479   0.03579  -0.0395   1.0000   0.0577
  -7.000  -0.5985   0.04185   0.03226  -0.0386   1.0000   0.0591
  -6.750  -0.5813   0.03903   0.02899  -0.0376   1.0000   0.0594
  -6.500  -0.5622   0.03656   0.02605  -0.0366   1.0000   0.0598
  -6.250  -0.5421   0.03410   0.02339  -0.0356   1.0000   0.0605
  -6.000  -0.5212   0.03208   0.02119  -0.0346   1.0000   0.0614
  -5.750  -0.5000   0.03039   0.01936  -0.0334   1.0000   0.0626
  -5.500  -0.4789   0.02895   0.01781  -0.0321   1.0000   0.0640
  -5.250  -0.4583   0.02770   0.01646  -0.0306   1.0000   0.0658
  -5.000  -0.4387   0.02662   0.01522  -0.0290   1.0000   0.0679
  -4.750  -0.4201   0.02570   0.01415  -0.0273   1.0000   0.0716
  -4.500  -0.4030   0.02475   0.01317  -0.0258   1.0000   0.0763
  -4.250  -0.3861   0.02388   0.01225  -0.0243   1.0000   0.0830
  -4.000  -0.3690   0.02300   0.01131  -0.0230   1.0000   0.0901
  -3.750  -0.3512   0.02207   0.01037  -0.0218   1.0000   0.1024
  -3.500  -0.3334   0.02075   0.00931  -0.0211   1.0000   0.1420
  -3.250  -0.3207   0.01853   0.00876  -0.0201   1.0000   0.4172
  -3.000  -0.3123   0.01831   0.00930  -0.0157   1.0000   0.6166
  -2.750  -0.3128   0.01879   0.01026  -0.0079   1.0000   0.7215
  -2.500  -0.3042   0.01940   0.01096  -0.0017   0.9941   0.7901
  -2.250  -0.2746   0.01961   0.01098  -0.0012   0.9843   0.8212
  -2.000  -0.2395   0.01965   0.01074  -0.0025   0.9751   0.8399
  -1.750  -0.2043   0.01965   0.01052  -0.0042   0.9656   0.8571
  -1.500  -0.1678   0.01964   0.01031  -0.0061   0.9565   0.8738
  -1.250  -0.1270   0.01966   0.01013  -0.0090   0.9486   0.8898
  -1.000  -0.0887   0.01963   0.00996  -0.0114   0.9392   0.9063
  -0.750  -0.0457   0.01962   0.00982  -0.0148   0.9311   0.9220
  -0.500   0.0018   0.01962   0.00969  -0.0191   0.9235   0.9368
  -0.250   0.0505   0.01960   0.00959  -0.0238   0.9156   0.9510
   0.000   0.1037   0.01957   0.00948  -0.0293   0.9087   0.9638
   0.250   0.1545   0.01952   0.00940  -0.0345   0.9003   0.9767
   0.500   0.2092   0.01943   0.00931  -0.0403   0.8934   0.9885
   0.750   0.2489   0.01942   0.00931  -0.0435   0.8829   1.0000
   1.000   0.2732   0.01949   0.00938  -0.0436   0.8709   1.0000
   1.250   0.2972   0.01961   0.00952  -0.0435   0.8592   1.0000
   1.500   0.3218   0.01976   0.00969  -0.0435   0.8482   1.0000
   1.750   0.3486   0.01991   0.00988  -0.0436   0.8378   1.0000
   2.000   0.3705   0.02018   0.01018  -0.0430   0.8253   1.0000
   2.250   0.3939   0.02046   0.01055  -0.0426   0.8132   1.0000
   2.500   0.4186   0.02075   0.01091  -0.0424   0.8013   1.0000
   2.750   0.4450   0.02101   0.01127  -0.0422   0.7899   1.0000
   3.000   0.4734   0.02121   0.01162  -0.0423   0.7790   1.0000
   3.250   0.4998   0.02146   0.01201  -0.0419   0.7666   1.0000
   3.500   0.5252   0.02170   0.01242  -0.0414   0.7530   1.0000
   3.750   0.5514   0.02188   0.01278  -0.0407   0.7385   1.0000
   4.000   0.5783   0.02195   0.01308  -0.0399   0.7227   1.0000
   4.250   0.6022   0.02197   0.01331  -0.0384   0.7015   1.0000
   4.500   0.6249   0.02103   0.01244  -0.0345   0.6561   1.0000
   4.750   0.6383   0.01995   0.01094  -0.0284   0.5493   1.0000
   5.000   0.6491   0.02045   0.01044  -0.0241   0.3581   1.0000
   5.250   0.6545   0.02288   0.01150  -0.0217   0.1610   1.0000
   5.500   0.6676   0.02492   0.01302  -0.0205   0.1009   1.0000
   5.750   0.6845   0.02644   0.01448  -0.0193   0.0842   1.0000
   6.250   0.7212   0.02916   0.01735  -0.0171   0.0716   1.0000
   6.500   0.7418   0.03049   0.01881  -0.0161   0.0679   1.0000
   6.750   0.7646   0.03191   0.02035  -0.0153   0.0651   1.0000
   7.000   0.7913   0.03364   0.02207  -0.0148   0.0629   1.0000
   7.250   0.8208   0.03533   0.02402  -0.0145   0.0612   1.0000
   7.500   0.8480   0.03719   0.02618  -0.0142   0.0584   1.0000
   7.750   0.8723   0.03916   0.02842  -0.0138   0.0554   1.0000
   8.000   0.8956   0.04147   0.03099  -0.0133   0.0536   1.0000
   8.250   0.9170   0.04412   0.03403  -0.0128   0.0527   1.0000
   8.500   0.9356   0.04706   0.03732  -0.0121   0.0520   1.0000
   8.750   0.9508   0.05026   0.04089  -0.0113   0.0515   1.0000
   9.000   0.9624   0.05374   0.04475  -0.0103   0.0511   1.0000
   9.250   0.9699   0.05745   0.04886  -0.0092   0.0509   1.0000
   9.500   0.9731   0.06133   0.05313  -0.0081   0.0508   1.0000
   9.750   0.9715   0.06534   0.05751  -0.0070   0.0508   1.0000
  10.000   0.9667   0.06940   0.06185  -0.0060   0.0505   1.0000
  10.500   0.9429   0.07769   0.07050  -0.0045   0.0498   1.0000
  10.750   0.9252   0.08190   0.07489  -0.0046   0.0498   1.0000
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