Airfoil Tools
Search 1636 airfoils Google+
You have 0 airfoils loaded.
Your Reynold number range is 50,000 to 1,000,000. (set)

USA 46 AIRFOIL (usa46-il) Xfoil prediction polar at RE=200,000 Ncrit=9


Details Polar file
Airfoil: USA 46 AIRFOIL (usa46-il)
Reynolds number: 200,000
Max Cl/Cd: 68.97 at α=3.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-usa46-il-200000.txt
Download as CSV file: xf-usa46-il-200000.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 =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.000  -0.4852   0.08619   0.08270  -0.0145   1.0000   0.0510
  -7.750  -0.4875   0.08273   0.07930  -0.0164   1.0000   0.0528
  -7.500  -0.4904   0.07783   0.07446  -0.0247   1.0000   0.0552
  -7.250  -0.4824   0.07074   0.06731  -0.0374   1.0000   0.0561
  -7.000  -0.4782   0.06333   0.05980  -0.0431   1.0000   0.0566
  -6.750  -0.4722   0.05935   0.05589  -0.0417   1.0000   0.0578
  -6.500  -0.4701   0.04059   0.03657  -0.0518   1.0000   0.0350
  -6.250  -0.4538   0.03529   0.03090  -0.0537   1.0000   0.0358
  -6.000  -0.4360   0.02791   0.02264  -0.0556   1.0000   0.0355
  -5.750  -0.4144   0.02407   0.01824  -0.0557   1.0000   0.0362
  -5.500  -0.3914   0.02158   0.01534  -0.0552   1.0000   0.0382
  -5.250  -0.3674   0.01967   0.01302  -0.0546   1.0000   0.0415
  -5.000  -0.3438   0.01775   0.01076  -0.0540   1.0000   0.0464
  -4.750  -0.3198   0.01666   0.00954  -0.0533   1.0000   0.0520
  -4.500  -0.2958   0.01537   0.00803  -0.0525   1.0000   0.0595
  -4.250  -0.2717   0.01486   0.00741  -0.0517   1.0000   0.0681
  -4.000  -0.2483   0.01396   0.00652  -0.0510   1.0000   0.0770
  -3.750  -0.2245   0.01327   0.00576  -0.0502   1.0000   0.0856
  -3.500  -0.2008   0.01292   0.00534  -0.0494   1.0000   0.0953
  -3.250  -0.1773   0.01237   0.00483  -0.0486   1.0000   0.1059
  -3.000  -0.1539   0.01203   0.00451  -0.0478   1.0000   0.1202
  -2.750  -0.1308   0.01177   0.00431  -0.0470   1.0000   0.1400
  -2.500  -0.1078   0.01157   0.00410  -0.0462   1.0000   0.1643
  -2.250  -0.0850   0.01128   0.00389  -0.0455   1.0000   0.1859
  -2.000  -0.0625   0.01115   0.00378  -0.0447   1.0000   0.2068
  -1.750  -0.0405   0.01096   0.00369  -0.0439   1.0000   0.2257
  -1.500  -0.0157   0.01083   0.00364  -0.0438   0.9991   0.2469
  -1.250   0.0289   0.01052   0.00355  -0.0476   0.9914   0.2876
  -1.000   0.0715   0.00954   0.00359  -0.0511   0.9835   0.5219
  -0.750   0.1036   0.00871   0.00370  -0.0511   0.9732   0.8121
  -0.500   0.1620   0.00824   0.00342  -0.0569   0.9684   1.0000
  -0.250   0.2080   0.00814   0.00322  -0.0605   0.9577   1.0000
   0.000   0.2518   0.00803   0.00304  -0.0635   0.9458   1.0000
   0.250   0.2923   0.00793   0.00290  -0.0658   0.9329   1.0000
   0.500   0.3307   0.00786   0.00278  -0.0676   0.9184   1.0000
   0.750   0.3667   0.00781   0.00269  -0.0688   0.9018   1.0000
   1.000   0.3976   0.00781   0.00265  -0.0690   0.8810   1.0000
   1.250   0.4265   0.00786   0.00265  -0.0686   0.8590   1.0000
   1.500   0.4531   0.00794   0.00267  -0.0679   0.8349   1.0000
   1.750   0.4793   0.00804   0.00271  -0.0671   0.8098   1.0000
   2.000   0.5051   0.00816   0.00276  -0.0662   0.7834   1.0000
   2.250   0.5306   0.00831   0.00284  -0.0653   0.7550   1.0000
   2.500   0.5559   0.00848   0.00293  -0.0643   0.7253   1.0000
   2.750   0.5811   0.00867   0.00304  -0.0634   0.6940   1.0000
   3.000   0.6061   0.00891   0.00317  -0.0624   0.6615   1.0000
   3.250   0.6307   0.00916   0.00335  -0.0614   0.6237   1.0000
   3.500   0.6545   0.00949   0.00351  -0.0603   0.5788   1.0000
   3.750   0.6772   0.00992   0.00368  -0.0590   0.5204   1.0000
   4.000   0.7004   0.01033   0.00390  -0.0579   0.4578   1.0000
   4.250   0.7230   0.01085   0.00416  -0.0569   0.3863   1.0000
   4.500   0.7449   0.01157   0.00454  -0.0559   0.3166   1.0000
   4.750   0.7591   0.01370   0.00540  -0.0543   0.0850   1.0000
   5.000   0.7796   0.01527   0.00676  -0.0529   0.0450   1.0000
   5.250   0.8014   0.01647   0.00804  -0.0517   0.0388   1.0000
   5.500   0.8245   0.01735   0.00898  -0.0507   0.0340   1.0000
   5.750   0.8454   0.01874   0.01041  -0.0495   0.0315   1.0000
   6.000   0.8653   0.02114   0.01282  -0.0480   0.0300   1.0000
   6.250   0.8895   0.02292   0.01469  -0.0470   0.0295   1.0000
   6.500   0.9142   0.02466   0.01660  -0.0461   0.0290   1.0000
   6.750   0.9384   0.02600   0.01817  -0.0451   0.0272   1.0000
   7.000   0.9617   0.02840   0.02086  -0.0440   0.0270   1.0000
   7.250   0.9830   0.03156   0.02443  -0.0426   0.0276   1.0000
   7.500   1.0002   0.03611   0.02937  -0.0411   0.0291   1.0000
  11.250   0.7625   0.12703   0.12403  -0.0577   0.0578   1.0000
  11.500   0.7632   0.13104   0.12805  -0.0588   0.0562   1.0000
<< Back to USA 46 AIRFOIL (usa46-il)

Polar data table (+)

Polar graphs


<< Back to USA 46 AIRFOIL (usa46-il)