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USA 46 AIRFOIL (usa46-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: USA 46 AIRFOIL (usa46-il)
Reynolds number: 50,000
Max Cl/Cd: 38.3 at α=5.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-usa46-il-50000.txt
Download as CSV file: xf-usa46-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: USA 46 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  -0.4826   0.10021   0.09332  -0.0029   1.0000   0.2181
  -7.750  -0.4800   0.09668   0.08986  -0.0034   1.0000   0.2242
  -7.500  -0.4858   0.09475   0.08804  -0.0043   1.0000   0.2338
  -7.250  -0.4740   0.09054   0.08384  -0.0033   1.0000   0.2454
  -7.000  -0.4708   0.08741   0.08074  -0.0030   1.0000   0.2588
  -6.750  -0.4700   0.08460   0.07801  -0.0034   1.0000   0.2751
  -6.500  -0.4729   0.08218   0.07569  -0.0055   1.0000   0.2914
  -6.250  -0.4617   0.07811   0.07165  -0.0035   1.0000   0.3082
  -6.000  -0.4509   0.07466   0.06823  -0.0015   1.0000   0.3282
  -5.750  -0.4462   0.07164   0.06527  -0.0006   1.0000   0.3524
  -5.500  -0.4455   0.06963   0.06332   0.0008   1.0000   0.3910
  -4.750  -0.3424   0.04019   0.03181  -0.0498   1.0000   0.1637
  -4.500  -0.3193   0.03646   0.02785  -0.0503   1.0000   0.1678
  -4.250  -0.2920   0.03327   0.02395  -0.0515   1.0000   0.1781
  -4.000  -0.2677   0.03039   0.02089  -0.0514   1.0000   0.1845
  -3.750  -0.2423   0.02815   0.01830  -0.0513   1.0000   0.1986
  -3.500  -0.2166   0.02610   0.01588  -0.0511   1.0000   0.2133
  -3.250  -0.1904   0.02418   0.01366  -0.0507   1.0000   0.2282
  -3.000  -0.1661   0.02272   0.01210  -0.0501   1.0000   0.2548
  -2.750  -0.1408   0.02124   0.01053  -0.0494   1.0000   0.2821
  -2.500  -0.1173   0.02004   0.00938  -0.0484   1.0000   0.3223
  -2.250  -0.0923   0.01894   0.00837  -0.0476   1.0000   0.3689
  -2.000  -0.0676   0.01794   0.00763  -0.0468   1.0000   0.4289
  -1.750  -0.0446   0.01683   0.00710  -0.0456   1.0000   0.5212
  -1.500  -0.0256   0.01460   0.00637  -0.0410   1.0000   1.0000
  -1.250   0.0024   0.01468   0.00580  -0.0415   1.0000   1.0000
  -1.000   0.0269   0.01479   0.00552  -0.0413   1.0000   1.0000
  -0.750   0.0505   0.01494   0.00539  -0.0409   1.0000   1.0000
  -0.500   0.0735   0.01511   0.00536  -0.0406   1.0000   1.0000
  -0.250   0.0959   0.01533   0.00539  -0.0401   1.0000   1.0000
   0.000   0.1179   0.01558   0.00549  -0.0397   1.0000   1.0000
   0.250   0.1393   0.01588   0.00567  -0.0393   1.0000   1.0000
   0.500   0.1601   0.01622   0.00593  -0.0388   1.0000   1.0000
   0.750   0.1802   0.01663   0.00627  -0.0385   1.0000   1.0000
   1.000   0.1996   0.01710   0.00669  -0.0381   1.0000   1.0000
   1.250   0.2183   0.01764   0.00721  -0.0378   1.0000   1.0000
   1.500   0.2363   0.01827   0.00783  -0.0376   1.0000   1.0000
   1.750   0.2535   0.01899   0.00854  -0.0375   1.0000   1.0000
   2.000   0.2701   0.01980   0.00937  -0.0376   1.0000   1.0000
   2.250   0.3270   0.02094   0.01067  -0.0450   0.9792   1.0000
   2.500   0.3933   0.02185   0.01176  -0.0534   0.9512   1.0000
   2.750   0.4576   0.02242   0.01258  -0.0606   0.9218   1.0000
   3.000   0.5241   0.02264   0.01317  -0.0674   0.8914   1.0000
   3.250   0.5887   0.02252   0.01342  -0.0727   0.8600   1.0000
   3.500   0.6374   0.02239   0.01366  -0.0746   0.8253   1.0000
   3.750   0.6791   0.02223   0.01377  -0.0747   0.7896   1.0000
   4.000   0.7108   0.02230   0.01405  -0.0732   0.7517   1.0000
   4.250   0.7402   0.02236   0.01428  -0.0709   0.7144   1.0000
   4.500   0.7662   0.02235   0.01443  -0.0679   0.6749   1.0000
   4.750   0.7876   0.02234   0.01449  -0.0641   0.6287   1.0000
   5.000   0.8040   0.02170   0.01369  -0.0584   0.5648   1.0000
   5.250   0.8127   0.02122   0.01289  -0.0523   0.4719   1.0000
   5.500   0.8112   0.02290   0.01322  -0.0459   0.2492   1.0000
   5.750   0.8247   0.02629   0.01536  -0.0436   0.1411   1.0000
   6.000   0.8483   0.02880   0.01772  -0.0420   0.1197   1.0000
   6.250   0.8752   0.03129   0.02016  -0.0409   0.1060   1.0000
   6.500   0.9027   0.03387   0.02295  -0.0400   0.0970   1.0000
   6.750   0.9302   0.03686   0.02613  -0.0391   0.0932   1.0000
   7.000   0.9559   0.04053   0.03005  -0.0383   0.0914   1.0000
   7.250   0.9781   0.04464   0.03455  -0.0373   0.0908   1.0000
   7.500   0.9954   0.04835   0.03885  -0.0358   0.0902   1.0000
   7.750   1.0088   0.05238   0.04355  -0.0344   0.0897   1.0000
   8.000   1.0206   0.05696   0.04858  -0.0333   0.0900   1.0000
   8.250   1.0189   0.06186   0.05441  -0.0318   0.0939   1.0000
   8.500   1.0098   0.06814   0.06127  -0.0315   0.0986   1.0000
   8.750   1.0114   0.07384   0.06714  -0.0314   0.1024   1.0000
   9.000   0.9864   0.08022   0.07393  -0.0328   0.1076   1.0000
   9.250   0.9544   0.08708   0.08086  -0.0354   0.1123   1.0000
   9.500   0.9254   0.09578   0.08958  -0.0417   0.1217   1.0000
   9.750   0.8992   0.10784   0.10155  -0.0517   0.1417   1.0000
  10.000   0.7681   0.11079   0.10483  -0.0486   0.1406   1.0000
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