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USA 49 AIRFOIL (usa49-il) Xfoil prediction polar at RE=100,000 Ncrit=9


Details Polar file
Airfoil: USA 49 AIRFOIL (usa49-il)
Reynolds number: 100,000
Max Cl/Cd: 43.92 at α=4.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-usa49-il-100000.txt
Download as CSV file: xf-usa49-il-100000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: USA 49 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
  ------ -------- --------- --------- -------- -------- --------
  -8.250  -0.5777   0.09811   0.09339   0.0028   1.0000   0.0938
  -8.000  -0.5917   0.09408   0.08944  -0.0080   1.0000   0.0962
  -7.750  -0.5939   0.08853   0.08390  -0.0139   1.0000   0.0975
  -7.500  -0.5759   0.08520   0.08063  -0.0064   1.0000   0.1010
  -7.250  -0.5697   0.08120   0.07663  -0.0091   1.0000   0.1058
  -7.000  -0.5748   0.07563   0.07083  -0.0211   1.0000   0.1113
  -6.750  -0.5582   0.07212   0.06750  -0.0156   1.0000   0.1158
  -6.500  -0.5547   0.06752   0.06256  -0.0235   1.0000   0.1253
  -6.250  -0.5388   0.06396   0.05920  -0.0196   1.0000   0.1312
  -6.000  -0.5284   0.05959   0.05466  -0.0225   1.0000   0.1416
  -5.750  -0.5155   0.05595   0.05078  -0.0244   1.0000   0.1542
  -5.500  -0.4992   0.05232   0.04722  -0.0233   1.0000   0.1596
  -5.250  -0.4838   0.04870   0.04346  -0.0239   1.0000   0.1719
  -5.000  -0.4672   0.04538   0.04000  -0.0239   1.0000   0.1857
  -4.750  -0.4257   0.03224   0.02483  -0.0262   1.0000   0.0661
  -4.500  -0.4000   0.02865   0.02023  -0.0244   1.0000   0.0576
  -4.250  -0.3770   0.02546   0.01674  -0.0234   1.0000   0.0563
  -4.000  -0.3524   0.02327   0.01417  -0.0223   1.0000   0.0568
  -3.750  -0.3283   0.02115   0.01186  -0.0213   1.0000   0.0601
  -3.500  -0.3035   0.01953   0.01014  -0.0203   1.0000   0.0625
  -3.250  -0.2789   0.01813   0.00863  -0.0191   1.0000   0.0660
  -3.000  -0.2550   0.01685   0.00733  -0.0179   1.0000   0.0718
  -2.750  -0.2318   0.01592   0.00653  -0.0169   1.0000   0.0888
  -2.500  -0.2097   0.01385   0.00523  -0.0158   1.0000   0.2096
  -2.250  -0.1166   0.01126   0.00511  -0.0250   1.0000   1.0000
  -2.000  -0.1005   0.01115   0.00477  -0.0231   1.0000   1.0000
  -1.750  -0.0849   0.01107   0.00450  -0.0211   1.0000   1.0000
  -1.500  -0.0692   0.01103   0.00428  -0.0191   1.0000   1.0000
  -1.250  -0.0525   0.01103   0.00410  -0.0173   1.0000   1.0000
  -1.000  -0.0348   0.01105   0.00398  -0.0156   1.0000   1.0000
  -0.750  -0.0160   0.01111   0.00391  -0.0142   1.0000   1.0000
  -0.500   0.0036   0.01119   0.00389  -0.0130   1.0000   1.0000
  -0.250   0.0237   0.01131   0.00392  -0.0120   1.0000   1.0000
   0.000   0.0443   0.01146   0.00398  -0.0111   1.0000   1.0000
   0.250   0.0650   0.01165   0.00412  -0.0104   1.0000   1.0000
   0.500   0.0856   0.01190   0.00433  -0.0098   1.0000   1.0000
   0.750   0.1058   0.01221   0.00463  -0.0094   1.0000   1.0000
   1.000   0.1428   0.01258   0.00500  -0.0123   0.9941   1.0000
   1.250   0.2037   0.01281   0.00527  -0.0196   0.9762   1.0000
   1.500   0.2656   0.01289   0.00544  -0.0266   0.9584   1.0000
   1.750   0.3230   0.01282   0.00550  -0.0324   0.9380   1.0000
   2.000   0.3714   0.01262   0.00540  -0.0357   0.9122   1.0000
   2.250   0.4051   0.01227   0.00510  -0.0351   0.8732   1.0000
   2.500   0.4295   0.01203   0.00486  -0.0326   0.8280   1.0000
   2.750   0.4529   0.01199   0.00476  -0.0305   0.7858   1.0000
   3.000   0.4764   0.01208   0.00478  -0.0287   0.7466   1.0000
   3.250   0.4993   0.01226   0.00484  -0.0268   0.7029   1.0000
   3.500   0.5221   0.01250   0.00500  -0.0251   0.6565   1.0000
   3.750   0.5450   0.01281   0.00516  -0.0235   0.6085   1.0000
   4.000   0.5678   0.01316   0.00536  -0.0220   0.5569   1.0000
   4.250   0.5909   0.01354   0.00562  -0.0208   0.5067   1.0000
   4.500   0.6135   0.01397   0.00588  -0.0195   0.4356   1.0000
   4.750   0.6339   0.01481   0.00624  -0.0180   0.3128   1.0000
   5.000   0.6519   0.01663   0.00709  -0.0168   0.1626   1.0000
   5.250   0.6740   0.01780   0.00802  -0.0159   0.1330   1.0000
   5.500   0.6971   0.01878   0.00900  -0.0150   0.1213   1.0000
   5.750   0.7201   0.01989   0.01014  -0.0140   0.1137   1.0000
   6.000   0.7428   0.02122   0.01144  -0.0131   0.1044   1.0000
   6.250   0.7666   0.02271   0.01288  -0.0123   0.0957   1.0000
   6.500   0.7909   0.02447   0.01475  -0.0116   0.0861   1.0000
   6.750   0.8156   0.02703   0.01740  -0.0108   0.0783   1.0000
   7.000   0.8396   0.02945   0.01999  -0.0101   0.0706   1.0000
   7.250   0.8620   0.03356   0.02442  -0.0092   0.0675   1.0000
   7.500   0.8817   0.03575   0.02737  -0.0076   0.0632   1.0000
   7.750   0.8994   0.03922   0.03141  -0.0061   0.0616   1.0000
   8.000   0.9134   0.04373   0.03653  -0.0045   0.0635   1.0000
   8.250   0.9253   0.04920   0.04236  -0.0033   0.0666   1.0000
   8.500   0.8617   0.04568   0.04047   0.0039   0.0869   1.0000
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