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USA 50 AIRFOIL (usa50-il) Xfoil prediction polar at RE=100,000 Ncrit=9


Details Polar file
Airfoil: USA 50 AIRFOIL (usa50-il)
Reynolds number: 100,000
Max Cl/Cd: 51.06 at α=4°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-usa50-il-100000.txt
Download as CSV file: xf-usa50-il-100000.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 =     0.100 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.250  -0.4091   0.10013   0.09540  -0.0205   1.0000   0.0768
  -9.000  -0.4215   0.09742   0.09277  -0.0235   1.0000   0.0775
  -8.750  -0.4333   0.09451   0.08994  -0.0262   1.0000   0.0778
  -8.500  -0.4088   0.08777   0.08317  -0.0215   1.0000   0.0808
  -8.250  -0.4052   0.08410   0.07953  -0.0210   1.0000   0.0837
  -8.000  -0.4081   0.08050   0.07597  -0.0216   1.0000   0.0865
  -7.750  -0.4167   0.07708   0.07263  -0.0230   1.0000   0.0889
  -7.500  -0.5073   0.08441   0.07978  -0.0221   1.0000   0.0809
  -7.250  -0.5034   0.08106   0.07646  -0.0219   1.0000   0.0836
  -7.000  -0.5025   0.07739   0.07281  -0.0238   1.0000   0.0869
  -6.750  -0.5084   0.07392   0.06911  -0.0318   1.0000   0.0912
  -6.500  -0.5023   0.06877   0.06405  -0.0306   1.0000   0.0931
  -6.250  -0.4930   0.06548   0.06084  -0.0282   1.0000   0.0966
  -6.000  -0.4907   0.06278   0.05761  -0.0326   1.0000   0.1054
  -5.750  -0.4791   0.05773   0.05287  -0.0299   1.0000   0.1084
  -5.500  -0.4705   0.05478   0.04955  -0.0311   1.0000   0.1196
  -5.250  -0.4582   0.05109   0.04599  -0.0290   1.0000   0.1240
  -5.000  -0.4469   0.04783   0.04257  -0.0286   1.0000   0.1362
  -4.750  -0.4344   0.04491   0.03954  -0.0276   1.0000   0.1502
  -4.500  -0.4211   0.04224   0.03680  -0.0263   1.0000   0.1677
  -4.000  -0.3645   0.03141   0.02415  -0.0248   1.0000   0.0737
  -3.750  -0.3435   0.02831   0.02056  -0.0231   1.0000   0.0701
  -3.500  -0.3187   0.02588   0.01736  -0.0209   1.0000   0.0632
  -3.250  -0.2956   0.02348   0.01459  -0.0194   1.0000   0.0625
  -3.000  -0.2712   0.02171   0.01240  -0.0179   1.0000   0.0638
  -2.750  -0.2476   0.01972   0.01035  -0.0169   1.0000   0.0698
  -2.500  -0.2240   0.01838   0.00887  -0.0157   1.0000   0.0820
  -2.250  -0.2009   0.01686   0.00741  -0.0143   1.0000   0.1006
  -2.000  -0.1805   0.01561   0.00636  -0.0128   1.0000   0.1470
  -1.750  -0.0908   0.01175   0.00547  -0.0242   1.0000   1.0000
  -1.500  -0.0719   0.01178   0.00520  -0.0225   1.0000   1.0000
  -1.250  -0.0529   0.01185   0.00497  -0.0209   1.0000   1.0000
  -1.000  -0.0339   0.01193   0.00486  -0.0194   1.0000   1.0000
  -0.750  -0.0149   0.01205   0.00480  -0.0180   1.0000   1.0000
  -0.500   0.0039   0.01219   0.00481  -0.0166   1.0000   1.0000
  -0.250   0.0225   0.01237   0.00487  -0.0153   1.0000   1.0000
   0.000   0.0405   0.01260   0.00497  -0.0140   1.0000   1.0000
   0.250   0.0578   0.01288   0.00518  -0.0127   1.0000   1.0000
   0.500   0.0740   0.01324   0.00548  -0.0114   1.0000   1.0000
   0.750   0.1020   0.01373   0.00593  -0.0128   0.9960   1.0000
   1.000   0.1628   0.01419   0.00637  -0.0203   0.9813   1.0000
   1.250   0.2236   0.01431   0.00650  -0.0274   0.9625   1.0000
   1.500   0.2843   0.01423   0.00648  -0.0341   0.9451   1.0000
   1.750   0.3366   0.01404   0.00640  -0.0389   0.9261   1.0000
   2.000   0.3916   0.01372   0.00620  -0.0441   0.9091   1.0000
   2.250   0.4490   0.01332   0.00595  -0.0495   0.8919   1.0000
   2.500   0.4966   0.01304   0.00581  -0.0529   0.8680   1.0000
   2.750   0.5411   0.01283   0.00578  -0.0556   0.8411   1.0000
   3.000   0.5804   0.01275   0.00580  -0.0572   0.8117   1.0000
   3.250   0.6135   0.01277   0.00587  -0.0574   0.7785   1.0000
   3.500   0.6420   0.01290   0.00599  -0.0566   0.7419   1.0000
   3.750   0.6625   0.01307   0.00604  -0.0537   0.6854   1.0000
   4.000   0.6816   0.01335   0.00625  -0.0511   0.6358   1.0000
   4.250   0.6984   0.01369   0.00639  -0.0479   0.5763   1.0000
   4.500   0.7117   0.01418   0.00659  -0.0443   0.4952   1.0000
   4.750   0.7187   0.01511   0.00676  -0.0397   0.3209   1.0000
   5.000   0.7165   0.01859   0.00849  -0.0348   0.0769   1.0000
   5.250   0.7315   0.02008   0.01004  -0.0320   0.0627   1.0000
   5.500   0.7478   0.02141   0.01145  -0.0297   0.0534   1.0000
   5.750   0.7662   0.02284   0.01295  -0.0276   0.0481   1.0000
   6.000   0.7879   0.02457   0.01473  -0.0260   0.0457   1.0000
   6.250   0.8131   0.02667   0.01692  -0.0249   0.0446   1.0000
   6.500   0.8398   0.02924   0.01970  -0.0239   0.0445   1.0000
   6.750   0.8642   0.03225   0.02314  -0.0224   0.0456   1.0000
   7.000   0.8843   0.03551   0.02687  -0.0203   0.0471   1.0000
   7.250   0.9006   0.03873   0.03055  -0.0181   0.0480   1.0000
   7.500   0.9131   0.04190   0.03419  -0.0155   0.0476   1.0000
   7.750   0.9235   0.04609   0.03876  -0.0131   0.0497   1.0000
   8.000   0.9287   0.05118   0.04464  -0.0094   0.0608   1.0000
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