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NACA 63-210 AIRFOIL (n63210-il) Xfoil prediction polar at RE=200,000 Ncrit=9


Details Polar file
Airfoil: NACA 63-210 AIRFOIL (n63210-il)
Reynolds number: 200,000
Max Cl/Cd: 59.42 at α=4°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-n63210-il-200000.txt
Download as CSV file: xf-n63210-il-200000.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 =     0.200 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -10.000  -0.4609   0.09075   0.08742  -0.0191   1.0000   0.0598
  -9.750  -0.5610   0.08983   0.08635  -0.0220   1.0000   0.0540
  -9.500  -0.5593   0.08597   0.08251  -0.0236   1.0000   0.0554
  -9.250  -0.5622   0.08121   0.07779  -0.0266   1.0000   0.0570
  -9.000  -0.5750   0.07419   0.07081  -0.0330   1.0000   0.0575
  -8.750  -0.5953   0.06830   0.06491  -0.0381   1.0000   0.0573
  -8.500  -0.6099   0.06385   0.06038  -0.0402   1.0000   0.0581
  -8.250  -0.6197   0.05948   0.05583  -0.0417   1.0000   0.0605
  -8.000  -0.6341   0.05884   0.05444  -0.0416   1.0000   0.0638
  -7.750  -0.6309   0.05044   0.04618  -0.0424   1.0000   0.0660
  -7.500  -0.6152   0.04816   0.04402  -0.0418   1.0000   0.0701
  -7.250  -0.6170   0.04531   0.04053  -0.0407   1.0000   0.0787
  -7.000  -0.6008   0.04199   0.03743  -0.0401   1.0000   0.0822
  -6.750  -0.5918   0.03256   0.02670  -0.0364   1.0000   0.0488
  -6.500  -0.5774   0.02972   0.02349  -0.0346   1.0000   0.0490
  -6.250  -0.5616   0.02744   0.02086  -0.0328   1.0000   0.0494
  -6.000  -0.5452   0.02605   0.01917  -0.0309   1.0000   0.0502
  -5.750  -0.5294   0.02418   0.01705  -0.0291   1.0000   0.0503
  -5.500  -0.5129   0.02226   0.01493  -0.0274   1.0000   0.0498
  -5.250  -0.4955   0.02084   0.01334  -0.0258   1.0000   0.0497
  -5.000  -0.4728   0.01973   0.01208  -0.0251   0.9991   0.0501
  -4.750  -0.4325   0.01882   0.01099  -0.0278   0.9945   0.0510
  -4.500  -0.3956   0.01735   0.00948  -0.0297   0.9897   0.0512
  -4.250  -0.3579   0.01595   0.00809  -0.0320   0.9850   0.0520
  -4.000  -0.3205   0.01486   0.00703  -0.0344   0.9795   0.0541
  -3.750  -0.2818   0.01410   0.00625  -0.0370   0.9735   0.0573
  -3.500  -0.2390   0.01349   0.00557  -0.0403   0.9696   0.0632
  -3.250  -0.2033   0.01263   0.00483  -0.0423   0.9619   0.0908
  -3.000  -0.1709   0.01023   0.00449  -0.0450   0.9568   0.5669
  -2.750  -0.1383   0.01009   0.00444  -0.0458   0.9480   0.6218
  -2.500  -0.1023   0.01003   0.00450  -0.0469   0.9421   0.6674
  -2.250  -0.0746   0.01010   0.00464  -0.0463   0.9316   0.7053
  -2.000  -0.0465   0.01019   0.00478  -0.0456   0.9224   0.7364
  -1.750  -0.0173   0.01020   0.00478  -0.0453   0.9141   0.7566
  -1.500   0.0071   0.01025   0.00485  -0.0438   0.9034   0.7759
  -1.250   0.0333   0.01024   0.00483  -0.0429   0.8939   0.7903
  -1.000   0.0609   0.01016   0.00470  -0.0425   0.8855   0.8011
  -0.750   0.0867   0.01015   0.00466  -0.0418   0.8750   0.8119
  -0.500   0.1118   0.01014   0.00463  -0.0409   0.8653   0.8216
  -0.250   0.1376   0.01014   0.00459  -0.0400   0.8566   0.8322
   0.000   0.1626   0.01017   0.00462  -0.0393   0.8457   0.8438
   0.250   0.1870   0.01020   0.00466  -0.0383   0.8357   0.8553
   0.500   0.2108   0.01022   0.00468  -0.0370   0.8265   0.8667
   0.750   0.2342   0.01024   0.00472  -0.0358   0.8162   0.8791
   1.000   0.2573   0.01025   0.00476  -0.0345   0.8061   0.8926
   1.250   0.2804   0.01024   0.00476  -0.0331   0.7975   0.9067
   1.500   0.3035   0.01021   0.00477  -0.0318   0.7874   0.9220
   1.750   0.3286   0.01017   0.00476  -0.0309   0.7774   0.9379
   2.000   0.3584   0.01011   0.00471  -0.0310   0.7686   0.9529
   2.250   0.3940   0.01004   0.00468  -0.0323   0.7570   0.9661
   2.500   0.4324   0.00984   0.00447  -0.0341   0.7358   0.9776
   2.750   0.4723   0.00960   0.00416  -0.0361   0.7067   0.9886
   3.000   0.5116   0.00947   0.00400  -0.0383   0.6766   1.0000
   3.250   0.5284   0.00948   0.00394  -0.0363   0.6498   1.0000
   3.500   0.5479   0.00953   0.00389  -0.0347   0.6130   1.0000
   3.750   0.5707   0.00968   0.00392  -0.0338   0.5698   1.0000
   4.000   0.5936   0.00999   0.00402  -0.0329   0.5020   1.0000
   4.500   0.6175   0.01405   0.00568  -0.0293   0.0648   1.0000
   5.000   0.6653   0.01558   0.00728  -0.0284   0.0515   1.0000
   5.250   0.6888   0.01635   0.00809  -0.0279   0.0491   1.0000
   5.500   0.7119   0.01722   0.00898  -0.0273   0.0472   1.0000
   5.750   0.7348   0.01820   0.00996  -0.0266   0.0459   1.0000
   6.000   0.7578   0.01936   0.01112  -0.0260   0.0449   1.0000
   6.250   0.7821   0.02069   0.01248  -0.0255   0.0443   1.0000
   6.500   0.8077   0.02219   0.01401  -0.0251   0.0440   1.0000
   6.750   0.8342   0.02373   0.01565  -0.0248   0.0440   1.0000
   7.000   0.8605   0.02583   0.01787  -0.0246   0.0435   1.0000
   7.250   0.8857   0.02860   0.02084  -0.0244   0.0429   1.0000
   7.500   0.9104   0.03045   0.02294  -0.0238   0.0432   1.0000
   7.750   0.9344   0.03208   0.02484  -0.0231   0.0438   1.0000
   8.000   0.9560   0.03462   0.02772  -0.0221   0.0441   1.0000
   8.250   0.9744   0.03788   0.03132  -0.0211   0.0439   1.0000
  13.250   0.7977   0.15763   0.15429  -0.0585   0.0502   1.0000
  13.500   0.8017   0.16192   0.15857  -0.0608   0.0472   1.0000
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