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USA 46 AIRFOIL (usa46-il) Xfoil prediction polar at RE=200,000 Ncrit=5


Details Polar file
Airfoil: USA 46 AIRFOIL (usa46-il)
Reynolds number: 200,000
Max Cl/Cd: 63.61 at α=3.5°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-usa46-il-200000-n5.txt
Download as CSV file: xf-usa46-il-200000-n5.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 =     0.200 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.750  -0.5155   0.08659   0.08310  -0.0136   1.0000   0.0118
  -8.500  -0.5185   0.08200   0.07856  -0.0157   1.0000   0.0117
  -8.250  -0.5232   0.07755   0.07417  -0.0178   1.0000   0.0116
  -8.000  -0.5298   0.07270   0.06940  -0.0205   1.0000   0.0116
  -7.750  -0.5322   0.06620   0.06295  -0.0267   1.0000   0.0116
  -7.500  -0.5312   0.05741   0.05415  -0.0364   1.0000   0.0114
  -7.250  -0.5326   0.04071   0.03714  -0.0500   1.0000   0.0109
  -7.000  -0.5229   0.03022   0.02576  -0.0555   1.0000   0.0109
  -6.750  -0.5045   0.02550   0.02035  -0.0566   1.0000   0.0113
  -6.500  -0.4833   0.02239   0.01663  -0.0567   1.0000   0.0119
  -6.250  -0.4601   0.02078   0.01465  -0.0563   1.0000   0.0128
  -6.000  -0.4382   0.01792   0.01134  -0.0560   1.0000   0.0149
  -5.750  -0.4145   0.01657   0.00976  -0.0554   1.0000   0.0167
  -5.500  -0.3903   0.01553   0.00853  -0.0548   1.0000   0.0193
  -5.250  -0.3659   0.01477   0.00758  -0.0541   1.0000   0.0234
  -5.000  -0.3415   0.01405   0.00681  -0.0536   1.0000   0.0310
  -4.750  -0.3171   0.01361   0.00628  -0.0530   1.0000   0.0406
  -4.500  -0.2928   0.01321   0.00582  -0.0524   1.0000   0.0494
  -4.250  -0.2687   0.01292   0.00546  -0.0517   1.0000   0.0589
  -4.000  -0.2447   0.01257   0.00505  -0.0511   1.0000   0.0660
  -3.750  -0.2207   0.01221   0.00459  -0.0504   1.0000   0.0707
  -3.500  -0.1932   0.01183   0.00419  -0.0505   0.9985   0.0796
  -3.250  -0.1581   0.01147   0.00382  -0.0522   0.9933   0.0902
  -3.000  -0.1231   0.01118   0.00356  -0.0538   0.9877   0.1062
  -2.750  -0.0889   0.01097   0.00334  -0.0552   0.9813   0.1265
  -2.500  -0.0541   0.01080   0.00316  -0.0567   0.9750   0.1448
  -2.250  -0.0201   0.01062   0.00299  -0.0580   0.9676   0.1624
  -2.000   0.0132   0.01044   0.00280  -0.0592   0.9596   0.1781
  -1.750   0.0486   0.01024   0.00264  -0.0608   0.9519   0.1951
  -1.500   0.0834   0.01000   0.00247  -0.0622   0.9405   0.2147
  -1.250   0.1192   0.00971   0.00229  -0.0638   0.9272   0.2466
  -1.000   0.1543   0.00926   0.00215  -0.0654   0.9123   0.3281
  -0.750   0.1859   0.00867   0.00210  -0.0662   0.8955   0.4929
  -0.500   0.2139   0.00825   0.00212  -0.0659   0.8758   0.6314
  -0.250   0.2411   0.00805   0.00212  -0.0651   0.8548   0.7296
   0.000   0.2672   0.00797   0.00207  -0.0643   0.8305   0.7835
   0.250   0.2935   0.00786   0.00200  -0.0634   0.8073   0.8257
   0.500   0.3292   0.00765   0.00191  -0.0645   0.7850   0.9069
   0.750   0.3622   0.00771   0.00188  -0.0654   0.7602   1.0000
   1.000   0.3887   0.00784   0.00189  -0.0648   0.7349   1.0000
   1.250   0.4149   0.00798   0.00192  -0.0643   0.7090   1.0000
   1.500   0.4411   0.00814   0.00197  -0.0637   0.6837   1.0000
   1.750   0.4672   0.00832   0.00206  -0.0631   0.6575   1.0000
   2.000   0.4932   0.00851   0.00215  -0.0626   0.6302   1.0000
   2.250   0.5189   0.00873   0.00226  -0.0620   0.6015   1.0000
   2.500   0.5446   0.00897   0.00242  -0.0614   0.5735   1.0000
   2.750   0.5704   0.00921   0.00259  -0.0608   0.5466   1.0000
   3.000   0.5959   0.00948   0.00278  -0.0603   0.5170   1.0000
   3.250   0.6212   0.00978   0.00301  -0.0597   0.4818   1.0000
   3.500   0.6456   0.01015   0.00322  -0.0589   0.4274   1.0000
   3.750   0.6672   0.01087   0.00347  -0.0579   0.3368   1.0000
   4.000   0.6912   0.01142   0.00383  -0.0573   0.2896   1.0000
   4.250   0.7150   0.01202   0.00422  -0.0567   0.2292   1.0000
   4.500   0.7340   0.01351   0.00491  -0.0558   0.0769   1.0000
   4.750   0.7559   0.01465   0.00575  -0.0549   0.0263   1.0000
   5.000   0.7801   0.01541   0.00658  -0.0541   0.0202   1.0000
   5.250   0.8045   0.01607   0.00739  -0.0534   0.0171   1.0000
   5.500   0.8279   0.01692   0.00839  -0.0526   0.0152   1.0000
   5.750   0.8500   0.01796   0.00960  -0.0516   0.0140   1.0000
   6.000   0.8702   0.01939   0.01115  -0.0503   0.0127   1.0000
   6.250   0.8891   0.02134   0.01322  -0.0489   0.0113   1.0000
   6.500   0.9116   0.02263   0.01464  -0.0479   0.0107   1.0000
   6.750   0.9336   0.02427   0.01646  -0.0468   0.0103   1.0000
   7.000   0.9554   0.02618   0.01859  -0.0456   0.0099   1.0000
   7.250   0.9766   0.02833   0.02101  -0.0445   0.0097   1.0000
   7.500   0.9963   0.03078   0.02379  -0.0432   0.0095   1.0000
   7.750   1.0144   0.03330   0.02672  -0.0418   0.0091   1.0000
   8.000   1.0312   0.03553   0.02926  -0.0405   0.0086   1.0000
   8.250   1.0472   0.03727   0.03117  -0.0395   0.0079   1.0000
   8.500   1.0598   0.03957   0.03371  -0.0383   0.0075   1.0000
   8.750   1.0678   0.04281   0.03726  -0.0367   0.0073   1.0000
   9.000   1.0717   0.04638   0.04119  -0.0349   0.0072   1.0000
   9.250   1.0710   0.05011   0.04526  -0.0330   0.0072   1.0000
   9.500   1.0654   0.05387   0.04934  -0.0311   0.0072   1.0000
   9.750   1.0537   0.05733   0.05305  -0.0288   0.0072   1.0000
  10.000   1.0390   0.06073   0.05667  -0.0271   0.0073   1.0000
  10.250   1.0210   0.06505   0.06119  -0.0274   0.0073   1.0000
  10.500   1.0033   0.07033   0.06666  -0.0298   0.0074   1.0000
  10.750   0.9819   0.07784   0.07438  -0.0351   0.0075   1.0000
  11.000   0.9491   0.09127   0.08806  -0.0456   0.0079   1.0000
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