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=100,000 Ncrit=9


Details Polar file
Airfoil: USA 46 AIRFOIL (usa46-il)
Reynolds number: 100,000
Max Cl/Cd: 52.92 at α=4.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-usa46-il-100000.txt
Download as CSV file: xf-usa46-il-100000.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 =     0.100 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.000  -0.4911   0.10620   0.10110  -0.0062   1.0000   0.0921
  -8.750  -0.5016   0.10424   0.09926  -0.0110   1.0000   0.0948
  -8.500  -0.5152   0.10232   0.09746  -0.0159   1.0000   0.0954
  -8.250  -0.4858   0.09569   0.09075  -0.0106   1.0000   0.0990
  -8.000  -0.4798   0.09256   0.08764  -0.0109   1.0000   0.1028
  -7.750  -0.4864   0.08994   0.08512  -0.0137   1.0000   0.1070
  -7.500  -0.5009   0.08652   0.08178  -0.0299   1.0000   0.1088
  -7.250  -0.4825   0.08221   0.07751  -0.0178   1.0000   0.1116
  -7.000  -0.4734   0.07894   0.07426  -0.0178   1.0000   0.1160
  -6.750  -0.4053   0.06351   0.05914  -0.0191   1.0000   0.1315
  -6.500  -0.4613   0.07030   0.06567  -0.0267   1.0000   0.1258
  -6.250  -0.4498   0.06523   0.06034  -0.0397   1.0000   0.1364
  -6.000  -0.4393   0.06269   0.05800  -0.0311   1.0000   0.1414
  -5.750  -0.4124   0.04390   0.03806  -0.0513   1.0000   0.0819
  -5.500  -0.3946   0.03925   0.03334  -0.0517   1.0000   0.0793
  -5.250  -0.3725   0.03307   0.02656  -0.0535   1.0000   0.0761
  -5.000  -0.3483   0.02852   0.02129  -0.0542   1.0000   0.0790
  -4.750  -0.3244   0.02549   0.01772  -0.0542   1.0000   0.0857
  -4.500  -0.2995   0.02338   0.01519  -0.0537   1.0000   0.0931
  -4.250  -0.2757   0.02214   0.01380  -0.0531   1.0000   0.1039
  -4.000  -0.2511   0.02067   0.01205  -0.0525   1.0000   0.1136
  -3.750  -0.2264   0.01953   0.01066  -0.0518   1.0000   0.1246
  -3.500  -0.2016   0.01853   0.00944  -0.0510   1.0000   0.1354
  -3.250  -0.1778   0.01774   0.00857  -0.0503   1.0000   0.1509
  -3.000  -0.1539   0.01679   0.00764  -0.0495   1.0000   0.1670
  -2.750  -0.1302   0.01595   0.00683  -0.0488   1.0000   0.1884
  -2.500  -0.1068   0.01523   0.00619  -0.0480   1.0000   0.2182
  -2.250  -0.0836   0.01455   0.00558  -0.0473   1.0000   0.2491
  -2.000  -0.0610   0.01396   0.00521  -0.0465   1.0000   0.2814
  -1.750  -0.0380   0.01348   0.00492  -0.0458   1.0000   0.3177
  -1.500  -0.0157   0.01297   0.00472  -0.0450   1.0000   0.3690
  -1.250   0.0051   0.01179   0.00481  -0.0435   1.0000   0.6250
  -1.000   0.0304   0.01092   0.00454  -0.0418   1.0000   1.0000
  -0.750   0.0532   0.01110   0.00449  -0.0413   1.0000   1.0000
  -0.500   0.0754   0.01132   0.00454  -0.0408   1.0000   1.0000
  -0.250   0.0968   0.01158   0.00467  -0.0403   1.0000   1.0000
   0.000   0.1176   0.01190   0.00487  -0.0398   1.0000   1.0000
   0.250   0.1376   0.01227   0.00516  -0.0393   1.0000   1.0000
   0.500   0.1999   0.01267   0.00548  -0.0467   0.9835   1.0000
   0.750   0.2549   0.01287   0.00565  -0.0524   0.9673   1.0000
   1.000   0.3063   0.01295   0.00574  -0.0571   0.9513   1.0000
   1.250   0.3577   0.01293   0.00575  -0.0616   0.9358   1.0000
   1.500   0.4086   0.01282   0.00573  -0.0658   0.9202   1.0000
   1.750   0.4553   0.01268   0.00566  -0.0688   0.9031   1.0000
   2.000   0.4953   0.01258   0.00563  -0.0704   0.8826   1.0000
   2.250   0.5304   0.01251   0.00565  -0.0709   0.8597   1.0000
   2.500   0.5601   0.01253   0.00572  -0.0703   0.8335   1.0000
   2.750   0.5876   0.01258   0.00580  -0.0692   0.8059   1.0000
   3.000   0.6137   0.01267   0.00591  -0.0678   0.7766   1.0000
   3.250   0.6391   0.01280   0.00606  -0.0663   0.7455   1.0000
   3.500   0.6631   0.01298   0.00623  -0.0646   0.7084   1.0000
   3.750   0.6868   0.01321   0.00641  -0.0628   0.6684   1.0000
   4.000   0.7098   0.01350   0.00660  -0.0610   0.6233   1.0000
   4.250   0.7315   0.01383   0.00680  -0.0589   0.5703   1.0000
   4.500   0.7515   0.01420   0.00692  -0.0566   0.5057   1.0000
   4.750   0.7708   0.01462   0.00707  -0.0546   0.4209   1.0000
   5.000   0.7878   0.01565   0.00754  -0.0525   0.2975   1.0000
   5.250   0.7966   0.01915   0.00944  -0.0498   0.0813   1.0000
   5.500   0.8164   0.02075   0.01098  -0.0483   0.0669   1.0000
   5.750   0.8371   0.02237   0.01270  -0.0467   0.0612   1.0000
   6.000   0.8596   0.02411   0.01442  -0.0454   0.0570   1.0000
   6.250   0.8830   0.02668   0.01689  -0.0445   0.0524   1.0000
   6.500   0.9090   0.02857   0.01904  -0.0435   0.0501   1.0000
   6.750   0.9347   0.03113   0.02188  -0.0425   0.0496   1.0000
   7.000   0.9587   0.03412   0.02525  -0.0412   0.0499   1.0000
   7.250   0.9800   0.03768   0.02924  -0.0399   0.0510   1.0000
   7.500   0.9984   0.04221   0.03417  -0.0387   0.0525   1.0000
   7.750   1.0160   0.04659   0.03890  -0.0376   0.0534   1.0000
   8.250   1.0344   0.05558   0.04958  -0.0326   0.0684   1.0000
  10.000   0.7487   0.11429   0.11010  -0.0552   0.1517   1.0000
  10.250   0.7475   0.11908   0.11488  -0.0564   0.1484   1.0000
<< Back to USA 46 AIRFOIL (usa46-il)

Polar data table (+)

Polar graphs


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