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USA 50 AIRFOIL (usa50-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: USA 50 AIRFOIL (usa50-il)
Reynolds number: 50,000
Max Cl/Cd: 36.45 at α=5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-usa50-il-50000.txt
Download as CSV file: xf-usa50-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: USA 50 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 =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.000  -0.4981   0.09927   0.09258  -0.0080   1.0000   0.2257
  -7.750  -0.4989   0.09641   0.08980  -0.0073   1.0000   0.2397
  -7.500  -0.5050   0.09397   0.08747  -0.0065   1.0000   0.2542
  -7.250  -0.4905   0.08994   0.08338  -0.0041   1.0000   0.2772
  -7.000  -0.4983   0.08770   0.08127  -0.0027   1.0000   0.2942
  -6.750  -0.4895   0.08418   0.07780  -0.0002   1.0000   0.3177
  -6.500  -0.5056   0.08263   0.07640   0.0001   1.0000   0.3373
  -6.250  -0.4831   0.07837   0.07213   0.0049   1.0000   0.3709
  -6.000  -0.4857   0.07608   0.06994   0.0080   1.0000   0.4047
  -5.750  -0.4761   0.07301   0.06692   0.0124   1.0000   0.4458
  -4.750  -0.3738   0.05795   0.05170   0.0260   1.0000   0.6787
  -4.500  -0.4134   0.04343   0.03556  -0.0285   1.0000   0.1561
  -4.250  -0.3884   0.03904   0.03038  -0.0284   1.0000   0.1291
  -4.000  -0.3647   0.03569   0.02627  -0.0273   1.0000   0.1172
  -3.750  -0.3424   0.03305   0.02305  -0.0259   1.0000   0.1181
  -3.500  -0.3205   0.03025   0.02006  -0.0248   1.0000   0.1228
  -3.250  -0.2954   0.02787   0.01724  -0.0235   1.0000   0.1237
  -3.000  -0.2692   0.02575   0.01469  -0.0222   1.0000   0.1289
  -2.750  -0.2433   0.02388   0.01249  -0.0210   1.0000   0.1467
  -2.500  -0.2138   0.02167   0.01037  -0.0204   1.0000   0.1779
  -2.250  -0.1253   0.01585   0.00783  -0.0288   1.0000   1.0000
  -2.000  -0.1050   0.01581   0.00707  -0.0270   1.0000   1.0000
  -1.750  -0.0857   0.01580   0.00660  -0.0254   1.0000   1.0000
  -1.500  -0.0663   0.01581   0.00618  -0.0240   1.0000   1.0000
  -1.250  -0.0468   0.01584   0.00591  -0.0226   1.0000   1.0000
  -1.000  -0.0273   0.01591   0.00571  -0.0212   1.0000   1.0000
  -0.750  -0.0078   0.01599   0.00558  -0.0199   1.0000   1.0000
  -0.500   0.0116   0.01611   0.00551  -0.0187   1.0000   1.0000
  -0.250   0.0309   0.01626   0.00550  -0.0174   1.0000   1.0000
   0.000   0.0498   0.01644   0.00552  -0.0162   1.0000   1.0000
   0.250   0.0684   0.01666   0.00563  -0.0150   1.0000   1.0000
   0.500   0.0863   0.01694   0.00582  -0.0139   1.0000   1.0000
   0.750   0.1036   0.01727   0.00609  -0.0127   1.0000   1.0000
   1.000   0.1199   0.01767   0.00646  -0.0116   1.0000   1.0000
   1.250   0.1350   0.01816   0.00693  -0.0105   1.0000   1.0000
   1.500   0.1491   0.01876   0.00751  -0.0096   1.0000   1.0000
   1.750   0.1621   0.01946   0.00821  -0.0087   1.0000   1.0000
   2.000   0.1746   0.02026   0.00901  -0.0080   1.0000   1.0000
   2.250   0.1871   0.02113   0.00989  -0.0075   1.0000   1.0000
   2.500   0.2431   0.02243   0.01130  -0.0153   0.9813   1.0000
   2.750   0.3082   0.02344   0.01253  -0.0240   0.9566   1.0000
   3.000   0.3674   0.02409   0.01340  -0.0310   0.9319   1.0000
   3.250   0.4229   0.02451   0.01410  -0.0368   0.9067   1.0000
   3.500   0.4783   0.02474   0.01473  -0.0420   0.8811   1.0000
   3.750   0.5371   0.02466   0.01506  -0.0472   0.8551   1.0000
   4.000   0.5972   0.02433   0.01523  -0.0519   0.8274   1.0000
   4.250   0.6538   0.02382   0.01531  -0.0551   0.7975   1.0000
   4.500   0.7011   0.02315   0.01512  -0.0555   0.7587   1.0000
   4.750   0.7359   0.02119   0.01324  -0.0494   0.6799   1.0000
   5.000   0.7450   0.02044   0.01219  -0.0412   0.5853   1.0000
   5.250   0.7447   0.02057   0.01192  -0.0331   0.4547   1.0000
   5.500   0.7337   0.02423   0.01313  -0.0258   0.1577   1.0000
   5.750   0.7486   0.02684   0.01535  -0.0235   0.1146   1.0000
   6.000   0.7719   0.02915   0.01764  -0.0218   0.1000   1.0000
   6.250   0.8031   0.03188   0.02031  -0.0213   0.0917   1.0000
   6.500   0.8350   0.03485   0.02360  -0.0207   0.0882   1.0000
   6.750   0.8595   0.03762   0.02681  -0.0193   0.0835   1.0000
   7.000   0.8822   0.04097   0.03028  -0.0184   0.0789   1.0000
   7.250   0.9013   0.04430   0.03412  -0.0165   0.0792   1.0000
   7.500   0.9124   0.04787   0.03855  -0.0136   0.0820   1.0000
   7.750   0.9210   0.05215   0.04343  -0.0111   0.0853   1.0000
   8.000   0.9287   0.05665   0.04830  -0.0091   0.0884   1.0000
   8.250   0.9321   0.06099   0.05315  -0.0069   0.0931   1.0000
   8.500   0.9190   0.06617   0.05890  -0.0048   0.0986   1.0000
   8.750   0.9235   0.07130   0.06422  -0.0039   0.1053   1.0000
   9.000   0.8910   0.07643   0.06972  -0.0031   0.1093   1.0000
   9.250   0.8874   0.08225   0.07564  -0.0032   0.1182   1.0000
   9.500   0.8523   0.08708   0.08052  -0.0042   0.1196   1.0000
   9.750   0.8190   0.09461   0.08800  -0.0105   0.1224   1.0000
  10.000   0.8000   0.10354   0.09684  -0.0177   0.1348   1.0000
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