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NACA 63A010 AIRFOIL (n63010a-il) Xfoil prediction polar at RE=100,000 Ncrit=9


Details Polar file
Airfoil: NACA 63A010 AIRFOIL (n63010a-il)
Reynolds number: 100,000
Max Cl/Cd: 37.68 at α=4°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-n63010a-il-100000.txt
Download as CSV file: xf-n63010a-il-100000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NACA 63A010 AIRFOIL                             
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.100 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.750  -0.5333   0.09589   0.09126  -0.0031   1.0000   0.1582
  -9.500  -0.5678   0.08998   0.08545  -0.0083   1.0000   0.1640
  -9.250  -0.7092   0.09003   0.08540  -0.0081   1.0000   0.1499
  -9.000  -0.6843   0.08783   0.08316  -0.0023   1.0000   0.1560
  -8.750  -0.7103   0.08144   0.07684  -0.0094   1.0000   0.1625
  -7.500  -0.7593   0.04430   0.03682  -0.0146   1.0000   0.0698
  -7.250  -0.7406   0.04136   0.03354  -0.0132   1.0000   0.0682
  -7.000  -0.7239   0.03769   0.02954  -0.0119   1.0000   0.0677
  -6.750  -0.7050   0.03416   0.02569  -0.0107   1.0000   0.0662
  -6.500  -0.6840   0.03122   0.02236  -0.0093   1.0000   0.0652
  -6.250  -0.6612   0.02874   0.01949  -0.0081   1.0000   0.0653
  -6.000  -0.6370   0.02666   0.01711  -0.0070   1.0000   0.0665
  -5.750  -0.6129   0.02480   0.01502  -0.0060   1.0000   0.0702
  -5.500  -0.5888   0.02298   0.01327  -0.0053   1.0000   0.0747
  -5.250  -0.5648   0.02153   0.01176  -0.0041   1.0000   0.0796
  -5.000  -0.5441   0.01997   0.01029  -0.0027   1.0000   0.0869
  -4.750  -0.5257   0.01873   0.00916  -0.0010   1.0000   0.1011
  -4.500  -0.5097   0.01737   0.00797   0.0011   1.0000   0.1256
  -4.250  -0.5073   0.01422   0.00655   0.0046   1.0000   0.3556
  -4.000  -0.5020   0.01332   0.00688   0.0102   1.0000   0.6258
  -3.750  -0.4854   0.01342   0.00706   0.0136   1.0000   0.6908
  -3.500  -0.4684   0.01359   0.00721   0.0168   1.0000   0.7354
  -3.250  -0.4516   0.01380   0.00744   0.0204   1.0000   0.7717
  -3.000  -0.4367   0.01406   0.00771   0.0245   1.0000   0.8070
  -2.750  -0.4225   0.01435   0.00800   0.0289   1.0000   0.8400
  -2.500  -0.4055   0.01458   0.00816   0.0324   1.0000   0.8682
  -2.250  -0.3785   0.01479   0.00828   0.0337   1.0000   0.8894
  -2.000  -0.3487   0.01485   0.00822   0.0335   1.0000   0.9091
  -1.750  -0.3135   0.01489   0.00813   0.0319   1.0000   0.9264
  -1.500  -0.2686   0.01495   0.00805   0.0282   1.0000   0.9404
  -1.250  -0.2145   0.01500   0.00798   0.0226   1.0000   0.9512
  -1.000  -0.1590   0.01500   0.00786   0.0164   1.0000   0.9616
  -0.750  -0.1070   0.01494   0.00774   0.0106   1.0000   0.9731
  -0.500  -0.0550   0.01485   0.00762   0.0047   1.0000   0.9847
  -0.250  -0.0016   0.01474   0.00748  -0.0018   1.0000   0.9957
   0.000   0.0000   0.01472   0.00746   0.0000   1.0000   1.0000
   0.250   0.0016   0.01474   0.00748   0.0018   0.9958   1.0000
   0.500   0.0550   0.01485   0.00762  -0.0047   0.9847   1.0000
   0.750   0.1069   0.01494   0.00774  -0.0106   0.9731   1.0000
   1.000   0.1590   0.01499   0.00786  -0.0164   0.9616   1.0000
   1.250   0.2145   0.01500   0.00798  -0.0226   0.9513   1.0000
   1.500   0.2686   0.01494   0.00804  -0.0282   0.9404   1.0000
   1.750   0.3135   0.01488   0.00813  -0.0319   0.9264   1.0000
   2.000   0.3486   0.01485   0.00822  -0.0335   0.9091   1.0000
   2.250   0.3784   0.01479   0.00827  -0.0336   0.8894   1.0000
   2.500   0.4054   0.01458   0.00816  -0.0324   0.8682   1.0000
   2.750   0.4224   0.01435   0.00799  -0.0288   0.8400   1.0000
   3.000   0.4366   0.01406   0.00771  -0.0244   0.8071   1.0000
   3.250   0.4514   0.01380   0.00744  -0.0203   0.7717   1.0000
   3.500   0.4682   0.01359   0.00721  -0.0168   0.7355   1.0000
   3.750   0.4853   0.01342   0.00706  -0.0135   0.6910   1.0000
   4.000   0.5019   0.01332   0.00688  -0.0101   0.6260   1.0000
   4.250   0.5073   0.01421   0.00655  -0.0046   0.3576   1.0000
   4.500   0.5095   0.01736   0.00796  -0.0011   0.1257   1.0000
   4.750   0.5256   0.01873   0.00916   0.0010   0.1011   1.0000
   5.000   0.5440   0.01997   0.01029   0.0027   0.0870   1.0000
   5.250   0.5647   0.02153   0.01175   0.0042   0.0796   1.0000
   5.500   0.5887   0.02297   0.01327   0.0053   0.0747   1.0000
   5.750   0.6128   0.02479   0.01501   0.0060   0.0702   1.0000
   6.000   0.6369   0.02666   0.01710   0.0070   0.0664   1.0000
   6.250   0.6611   0.02874   0.01949   0.0081   0.0652   1.0000
   6.500   0.6840   0.03122   0.02236   0.0093   0.0652   1.0000
   6.750   0.7050   0.03417   0.02569   0.0107   0.0662   1.0000
   7.000   0.7239   0.03770   0.02954   0.0119   0.0677   1.0000
   7.250   0.7406   0.04137   0.03356   0.0132   0.0682   1.0000
   7.500   0.7594   0.04430   0.03683   0.0146   0.0698   1.0000
   9.000   0.6854   0.08781   0.08313   0.0023   0.1559   1.0000
   9.250   0.7087   0.09015   0.08552   0.0077   0.1498   1.0000
   9.500   0.6706   0.09782   0.09303  -0.0036   0.1433   1.0000
   9.750   0.7220   0.09830   0.09363   0.0086   0.1367   1.0000
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