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USA 46 AIRFOIL (usa46-il) Xfoil prediction polar at RE=50,000 Ncrit=5


Details Polar file
Airfoil: USA 46 AIRFOIL (usa46-il)
Reynolds number: 50,000
Max Cl/Cd: 38.34 at α=5°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-usa46-il-50000-n5.txt
Download as CSV file: xf-usa46-il-50000-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 =     0.050 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.250  -0.4966   0.09234   0.08550  -0.0148   1.0000   0.0467
  -8.000  -0.4985   0.08809   0.08134  -0.0171   1.0000   0.0461
  -7.750  -0.4999   0.08334   0.07670  -0.0205   1.0000   0.0454
  -7.500  -0.4984   0.07778   0.07121  -0.0257   1.0000   0.0447
  -6.750  -0.4764   0.06285   0.05625  -0.0375   1.0000   0.0470
  -6.500  -0.4654   0.05751   0.05077  -0.0415   1.0000   0.0490
  -6.250  -0.4521   0.05140   0.04440  -0.0459   1.0000   0.0509
  -6.000  -0.4357   0.04490   0.03742  -0.0500   1.0000   0.0525
  -5.750  -0.4154   0.03839   0.03012  -0.0531   1.0000   0.0547
  -5.500  -0.3947   0.03591   0.02744  -0.0535   1.0000   0.0613
  -5.250  -0.3701   0.03167   0.02228  -0.0546   1.0000   0.0694
  -5.000  -0.3475   0.03008   0.02050  -0.0544   1.0000   0.0803
  -4.750  -0.3236   0.02822   0.01833  -0.0541   1.0000   0.0913
  -4.500  -0.2986   0.02611   0.01578  -0.0537   1.0000   0.1000
  -4.250  -0.2747   0.02527   0.01473  -0.0533   1.0000   0.1146
  -4.000  -0.2495   0.02386   0.01292  -0.0527   1.0000   0.1250
  -3.750  -0.2254   0.02276   0.01167  -0.0519   1.0000   0.1345
  -3.500  -0.2008   0.02190   0.01062  -0.0513   1.0000   0.1502
  -3.250  -0.1768   0.02105   0.00959  -0.0505   1.0000   0.1669
  -3.000  -0.1531   0.02023   0.00876  -0.0498   1.0000   0.1854
  -2.750  -0.1285   0.01955   0.00797  -0.0493   1.0000   0.2101
  -2.500  -0.1043   0.01888   0.00730  -0.0488   1.0000   0.2358
  -2.250  -0.0801   0.01830   0.00677  -0.0483   1.0000   0.2677
  -2.000  -0.0560   0.01769   0.00634  -0.0479   1.0000   0.3060
  -1.750  -0.0320   0.01703   0.00597  -0.0476   1.0000   0.3662
  -1.500  -0.0105   0.01612   0.00583  -0.0465   1.0000   0.5061
  -1.250   0.0029   0.01533   0.00581  -0.0426   1.0000   0.7442
  -1.000   0.0298   0.01479   0.00542  -0.0420   1.0000   1.0000
  -0.750   0.0534   0.01493   0.00528  -0.0417   1.0000   1.0000
  -0.500   0.0766   0.01511   0.00524  -0.0413   1.0000   1.0000
  -0.250   0.0992   0.01533   0.00528  -0.0409   1.0000   1.0000
   0.000   0.1213   0.01558   0.00538  -0.0405   1.0000   1.0000
   0.250   0.1505   0.01590   0.00557  -0.0416   0.9950   1.0000
   0.500   0.1937   0.01623   0.00582  -0.0453   0.9803   1.0000
   0.750   0.2386   0.01653   0.00606  -0.0492   0.9644   1.0000
   1.000   0.2852   0.01669   0.00621  -0.0530   0.9434   1.0000
   1.250   0.3352   0.01676   0.00632  -0.0571   0.9220   1.0000
   1.500   0.3799   0.01679   0.00640  -0.0599   0.9001   1.0000
   1.750   0.4194   0.01687   0.00655  -0.0618   0.8794   1.0000
   2.000   0.4562   0.01694   0.00674  -0.0630   0.8574   1.0000
   2.250   0.4896   0.01705   0.00693  -0.0635   0.8334   1.0000
   2.500   0.5218   0.01716   0.00714  -0.0637   0.8084   1.0000
   2.750   0.5523   0.01728   0.00740  -0.0634   0.7823   1.0000
   3.000   0.5812   0.01744   0.00765  -0.0629   0.7545   1.0000
   3.250   0.6087   0.01764   0.00795  -0.0620   0.7252   1.0000
   3.500   0.6354   0.01788   0.00834  -0.0610   0.6949   1.0000
   3.750   0.6614   0.01816   0.00873  -0.0599   0.6631   1.0000
   4.000   0.6871   0.01849   0.00917  -0.0587   0.6312   1.0000
   4.250   0.7124   0.01889   0.00966  -0.0575   0.5990   1.0000
   4.500   0.7370   0.01935   0.01030  -0.0561   0.5625   1.0000
   4.750   0.7605   0.01984   0.01088  -0.0546   0.5208   1.0000
   5.000   0.7783   0.02030   0.01120  -0.0519   0.4467   1.0000
   5.250   0.7878   0.02155   0.01135  -0.0488   0.2885   1.0000
   5.500   0.8014   0.02425   0.01284  -0.0474   0.1009   1.0000
   5.750   0.8193   0.02673   0.01495  -0.0463   0.0624   1.0000
   6.000   0.8387   0.02863   0.01708  -0.0449   0.0528   1.0000
   6.250   0.8578   0.03055   0.01920  -0.0435   0.0469   1.0000
   6.500   0.8772   0.03242   0.02121  -0.0422   0.0411   1.0000
   6.750   0.8976   0.03468   0.02359  -0.0409   0.0384   1.0000
   7.000   0.9214   0.03699   0.02622  -0.0396   0.0366   1.0000
   7.250   0.9444   0.03954   0.02910  -0.0384   0.0345   1.0000
   7.500   0.9645   0.04215   0.03195  -0.0375   0.0321   1.0000
   7.750   0.9820   0.04542   0.03538  -0.0367   0.0304   1.0000
   8.000   0.9978   0.04884   0.03936  -0.0353   0.0299   1.0000
   8.250   1.0098   0.05265   0.04369  -0.0340   0.0296   1.0000
   8.500   1.0175   0.05666   0.04831  -0.0327   0.0295   1.0000
   8.750   1.0208   0.06084   0.05294  -0.0314   0.0295   1.0000
   9.000   1.0200   0.06508   0.05756  -0.0304   0.0296   1.0000
   9.250   1.0149   0.06939   0.06218  -0.0295   0.0297   1.0000
   9.500   1.0054   0.07370   0.06675  -0.0290   0.0298   1.0000
   9.750   0.9913   0.07781   0.07104  -0.0285   0.0299   1.0000
  10.000   0.9764   0.08226   0.07563  -0.0293   0.0300   1.0000
  10.250   0.9607   0.08746   0.08094  -0.0316   0.0302   1.0000
  10.500   0.9469   0.09322   0.08678  -0.0350   0.0304   1.0000
  10.750   0.9349   0.09954   0.09315  -0.0390   0.0305   1.0000
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