Airfoil Tools
Search 1636 airfoils Google+
You have 0 airfoils loaded.
Your Reynold number range is 50,000 to 1,000,000. (set)

NACA 63A010 AIRFOIL (n63010a-il) Xfoil prediction polar at RE=50,000 Ncrit=5


Details Polar file
Airfoil: NACA 63A010 AIRFOIL (n63010a-il)
Reynolds number: 50,000
Max Cl/Cd: 26.38 at α=4.25°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-n63010a-il-50000-n5.txt
Download as CSV file: xf-n63010a-il-50000-n5.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 =     0.050 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -10.500  -0.7052   0.09375   0.08675  -0.0112   1.0000   0.0473
 -10.250  -0.7169   0.08745   0.08044  -0.0154   1.0000   0.0468
 -10.000  -0.7318   0.08164   0.07458  -0.0190   1.0000   0.0464
  -9.750  -0.7481   0.07658   0.06944  -0.0212   1.0000   0.0460
  -9.500  -0.7639   0.07222   0.06496  -0.0217   1.0000   0.0457
  -9.250  -0.7754   0.06787   0.06041  -0.0217   1.0000   0.0455
  -9.000  -0.7824   0.06362   0.05589  -0.0212   1.0000   0.0454
  -8.750  -0.7851   0.05947   0.05142  -0.0205   1.0000   0.0453
  -8.500  -0.7838   0.05547   0.04704  -0.0194   1.0000   0.0454
  -8.250  -0.7786   0.05166   0.04275  -0.0182   1.0000   0.0456
  -8.000  -0.7700   0.04817   0.03875  -0.0169   1.0000   0.0466
  -7.750  -0.7585   0.04512   0.03507  -0.0154   1.0000   0.0483
  -7.500  -0.7430   0.04200   0.03170  -0.0145   1.0000   0.0502
  -7.250  -0.7242   0.03937   0.02882  -0.0136   1.0000   0.0518
  -7.000  -0.7029   0.03687   0.02600  -0.0127   1.0000   0.0533
  -6.750  -0.6796   0.03459   0.02343  -0.0119   1.0000   0.0553
  -6.500  -0.6545   0.03256   0.02112  -0.0111   1.0000   0.0581
  -6.250  -0.6300   0.03081   0.01915  -0.0103   1.0000   0.0630
  -6.000  -0.6069   0.02931   0.01762  -0.0094   1.0000   0.0690
  -5.750  -0.5834   0.02790   0.01599  -0.0082   1.0000   0.0746
  -5.500  -0.5641   0.02654   0.01465  -0.0069   1.0000   0.0847
  -5.250  -0.5455   0.02520   0.01330  -0.0056   1.0000   0.0981
  -5.000  -0.5284   0.02373   0.01191  -0.0041   1.0000   0.1192
  -4.750  -0.5141   0.02186   0.01052  -0.0026   1.0000   0.1757
  -4.500  -0.5083   0.01946   0.00962   0.0003   1.0000   0.3869
  -4.250  -0.4973   0.01885   0.00975   0.0045   1.0000   0.5597
  -4.000  -0.4824   0.01885   0.00992   0.0085   1.0000   0.6455
  -3.750  -0.4678   0.01918   0.01035   0.0132   1.0000   0.7156
  -3.500  -0.4497   0.01956   0.01067   0.0174   1.0000   0.7643
  -3.250  -0.4272   0.01969   0.01066   0.0199   1.0000   0.7949
  -3.000  -0.4042   0.01961   0.01041   0.0214   1.0000   0.8178
  -2.750  -0.3783   0.01949   0.01012   0.0222   1.0000   0.8363
  -2.500  -0.3507   0.01935   0.00978   0.0223   1.0000   0.8530
  -2.250  -0.3222   0.01920   0.00948   0.0222   1.0000   0.8692
  -2.000  -0.2922   0.01907   0.00920   0.0215   1.0000   0.8852
  -1.750  -0.2604   0.01894   0.00892   0.0204   1.0000   0.9009
  -1.500  -0.2263   0.01883   0.00869   0.0186   1.0000   0.9163
  -1.250  -0.1899   0.01874   0.00850   0.0163   1.0000   0.9314
  -1.000  -0.1509   0.01866   0.00831   0.0133   1.0000   0.9458
  -0.750  -0.1094   0.01857   0.00816   0.0096   1.0000   0.9596
  -0.500  -0.0688   0.01850   0.00804   0.0059   1.0000   0.9742
  -0.250  -0.0270   0.01842   0.00795   0.0018   1.0000   0.9885
   0.000   0.0000   0.01838   0.00791   0.0000   1.0000   1.0000
   0.250   0.0270   0.01842   0.00795  -0.0018   0.9885   1.0000
   0.500   0.0688   0.01850   0.00804  -0.0059   0.9742   1.0000
   0.750   0.1094   0.01857   0.00816  -0.0096   0.9596   1.0000
   1.000   0.1508   0.01865   0.00831  -0.0133   0.9458   1.0000
   1.250   0.1899   0.01874   0.00849  -0.0163   0.9314   1.0000
   1.500   0.2263   0.01883   0.00869  -0.0186   0.9163   1.0000
   1.750   0.2603   0.01894   0.00892  -0.0204   0.9009   1.0000
   2.000   0.2922   0.01906   0.00920  -0.0215   0.8852   1.0000
   2.250   0.3222   0.01920   0.00948  -0.0221   0.8692   1.0000
   2.500   0.3507   0.01935   0.00978  -0.0223   0.8530   1.0000
   2.750   0.3782   0.01949   0.01012  -0.0222   0.8363   1.0000
   3.000   0.4042   0.01961   0.01041  -0.0214   0.8178   1.0000
   3.250   0.4271   0.01969   0.01066  -0.0199   0.7949   1.0000
   3.500   0.4496   0.01956   0.01067  -0.0174   0.7643   1.0000
   3.750   0.4678   0.01918   0.01035  -0.0132   0.7157   1.0000
   4.000   0.4824   0.01885   0.00992  -0.0084   0.6456   1.0000
   4.250   0.4972   0.01885   0.00975  -0.0045   0.5597   1.0000
   4.500   0.5083   0.01946   0.00962  -0.0003   0.3869   1.0000
   4.750   0.5141   0.02185   0.01052   0.0026   0.1757   1.0000
   5.000   0.5284   0.02373   0.01191   0.0041   0.1192   1.0000
   5.250   0.5456   0.02520   0.01330   0.0056   0.0981   1.0000
   5.500   0.5641   0.02654   0.01465   0.0069   0.0847   1.0000
   5.750   0.5834   0.02790   0.01599   0.0082   0.0746   1.0000
   6.000   0.6070   0.02931   0.01762   0.0094   0.0690   1.0000
   6.250   0.6300   0.03081   0.01915   0.0103   0.0630   1.0000
   6.500   0.6545   0.03256   0.02112   0.0111   0.0581   1.0000
   6.750   0.6796   0.03459   0.02343   0.0119   0.0553   1.0000
   7.000   0.7030   0.03687   0.02600   0.0127   0.0533   1.0000
   7.250   0.7242   0.03937   0.02882   0.0136   0.0518   1.0000
   7.500   0.7431   0.04200   0.03170   0.0145   0.0502   1.0000
   7.750   0.7585   0.04512   0.03507   0.0154   0.0483   1.0000
   8.000   0.7701   0.04817   0.03875   0.0169   0.0466   1.0000
   8.250   0.7787   0.05166   0.04275   0.0182   0.0456   1.0000
   8.500   0.7839   0.05548   0.04705   0.0194   0.0454   1.0000
   8.750   0.7852   0.05948   0.05143   0.0204   0.0453   1.0000
   9.000   0.7825   0.06362   0.05590   0.0212   0.0454   1.0000
   9.250   0.7756   0.06788   0.06042   0.0216   0.0455   1.0000
   9.500   0.7642   0.07224   0.06498   0.0216   0.0457   1.0000
   9.750   0.7484   0.07661   0.06947   0.0212   0.0460   1.0000
  10.000   0.7322   0.08167   0.07462   0.0189   0.0464   1.0000
  10.250   0.7174   0.08750   0.08049   0.0153   0.0468   1.0000
  10.500   0.7057   0.09381   0.08681   0.0111   0.0473   1.0000
<< Back to NACA 63A010 AIRFOIL (n63010a-il)

Polar data table (+)

Polar graphs


<< Back to NACA 63A010 AIRFOIL (n63010a-il)