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

FX 66-H-60 (fx66h60-il) Xfoil prediction polar at RE=50,000 Ncrit=5


Details Polar file
Airfoil: FX 66-H-60 (fx66h60-il)
Reynolds number: 50,000
Max Cl/Cd: 29.85 at α=6.5°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-fx66h60-il-50000-n5.txt
Download as CSV file: xf-fx66h60-il-50000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: FX 66-H-60                                      
  
 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
  ------ -------- --------- --------- -------- -------- --------
  -8.750  -0.6241   0.11515   0.10874   0.0343   1.0000   0.0800
  -8.500  -0.6277   0.11226   0.10593   0.0306   1.0000   0.0823
  -8.250  -0.6341   0.10936   0.10313   0.0261   1.0000   0.0831
  -8.000  -0.6190   0.10375   0.09755   0.0289   1.0000   0.0888
  -7.750  -0.6191   0.10024   0.09411   0.0267   1.0000   0.0929
  -7.500  -0.6262   0.09738   0.09130   0.0214   1.0000   0.0960
  -7.000  -0.6128   0.08872   0.08271   0.0204   1.0000   0.1059
  -6.500  -0.6013   0.08071   0.07472   0.0168   1.0000   0.1170
  -6.250  -0.5967   0.07708   0.07099   0.0134   1.0000   0.1255
  -6.000  -0.5844   0.07291   0.06684   0.0138   1.0000   0.1328
  -5.750  -0.5742   0.06894   0.06282   0.0129   1.0000   0.1434
  -5.500  -0.5218   0.05348   0.04791   0.0111   1.0000   0.1680
  -5.000  -0.5041   0.05518   0.04799   0.0074   1.0000   0.0681
  -4.750  -0.4820   0.05059   0.04314   0.0075   1.0000   0.0527
  -4.250  -0.4343   0.04349   0.03511   0.0085   1.0000   0.0424
  -4.000  -0.4123   0.04010   0.03146   0.0090   1.0000   0.0411
  -3.750  -0.3885   0.03708   0.02807   0.0096   1.0000   0.0398
  -3.500  -0.3633   0.03428   0.02475   0.0104   1.0000   0.0388
  -3.250  -0.3370   0.03172   0.02179   0.0111   1.0000   0.0381
  -3.000  -0.3098   0.02938   0.01907   0.0119   1.0000   0.0377
  -2.750  -0.2818   0.02730   0.01663   0.0126   1.0000   0.0385
  -2.500  -0.2531   0.02571   0.01463   0.0134   1.0000   0.0425
  -2.250  -0.2248   0.02383   0.01265   0.0137   1.0000   0.0470
  -2.000  -0.1938   0.02225   0.01092   0.0141   1.0000   0.0500
  -1.750  -0.1633   0.02089   0.00935   0.0145   1.0000   0.0541
  -1.500  -0.1376   0.01968   0.00804   0.0153   1.0000   0.0606
  -1.250  -0.0388   0.01461   0.00598   0.0040   1.0000   1.0000
  -1.000  -0.0142   0.01451   0.00548   0.0046   1.0000   1.0000
  -0.750   0.0104   0.01443   0.00518   0.0049   1.0000   1.0000
  -0.500   0.0525   0.01467   0.00496   0.0020   0.8261   1.0000
  -0.250   0.0796   0.01511   0.00478   0.0028   0.7385   1.0000
   0.000   0.1018   0.01547   0.00469   0.0043   0.6846   1.0000
   0.250   0.1240   0.01579   0.00463   0.0057   0.6434   1.0000
   0.500   0.1468   0.01609   0.00461   0.0069   0.6098   1.0000
   0.750   0.1700   0.01637   0.00463   0.0079   0.5815   1.0000
   1.000   0.1936   0.01666   0.00469   0.0088   0.5564   1.0000
   1.250   0.2176   0.01694   0.00479   0.0096   0.5341   1.0000
   1.500   0.2416   0.01725   0.00492   0.0103   0.5148   1.0000
   1.750   0.2659   0.01755   0.00511   0.0110   0.4966   1.0000
   2.000   0.2903   0.01787   0.00534   0.0116   0.4799   1.0000
   2.250   0.3148   0.01820   0.00562   0.0122   0.4650   1.0000
   2.500   0.3394   0.01855   0.00601   0.0128   0.4511   1.0000
   2.750   0.3640   0.01892   0.00637   0.0134   0.4381   1.0000
   3.000   0.3886   0.01931   0.00679   0.0140   0.4261   1.0000
   3.250   0.4127   0.01973   0.00725   0.0146   0.4151   1.0000
   3.500   0.4367   0.02018   0.00772   0.0153   0.4050   1.0000
   3.750   0.4613   0.02066   0.00831   0.0158   0.3948   1.0000
   4.000   0.4862   0.02119   0.00912   0.0163   0.3848   1.0000
   4.250   0.5107   0.02174   0.00981   0.0168   0.3761   1.0000
   4.500   0.5350   0.02232   0.01054   0.0174   0.3677   1.0000
   4.750   0.5594   0.02299   0.01146   0.0178   0.3588   1.0000
   5.000   0.5836   0.02364   0.01227   0.0184   0.3516   1.0000
   5.250   0.6079   0.02442   0.01339   0.0188   0.3432   1.0000
   5.500   0.6321   0.02522   0.01449   0.0193   0.3355   1.0000
   5.750   0.6562   0.02604   0.01583   0.0199   0.3280   1.0000
   6.000   0.6802   0.02704   0.01731   0.0203   0.3197   1.0000
   6.250   0.7028   0.02664   0.01711   0.0219   0.2982   1.0000
   6.500   0.7215   0.02417   0.01467   0.0237   0.2030   1.0000
   6.750   0.7360   0.02701   0.01648   0.0237   0.0662   1.0000
   7.000   0.7515   0.02994   0.01945   0.0239   0.0482   1.0000
   7.250   0.7654   0.03267   0.02233   0.0240   0.0400   1.0000
   7.500   0.7807   0.03487   0.02486   0.0248   0.0368   1.0000
   7.750   0.7947   0.03719   0.02742   0.0258   0.0342   1.0000
   8.000   0.8083   0.03957   0.03000   0.0269   0.0325   1.0000
   8.250   0.8217   0.04224   0.03285   0.0281   0.0311   1.0000
   8.500   0.8332   0.04571   0.03658   0.0292   0.0295   1.0000
   8.750   0.8439   0.04864   0.04001   0.0294   0.0282   1.0000
   9.000   0.8496   0.05208   0.04389   0.0292   0.0269   1.0000
   9.250   0.8501   0.05590   0.04808   0.0283   0.0259   1.0000
   9.500   0.8459   0.06002   0.05248   0.0268   0.0257   1.0000
   9.750   0.8352   0.06452   0.05715   0.0241   0.0256   1.0000
  10.000   0.8246   0.06992   0.06270   0.0200   0.0260   1.0000
  10.250   0.8142   0.07576   0.06867   0.0157   0.0264   1.0000
  10.500   0.8026   0.08224   0.07524   0.0112   0.0268   1.0000
  10.750   0.7915   0.08887   0.08193   0.0068   0.0272   1.0000
<< Back to FX 66-H-60 (fx66h60-il)

Polar data table (+)

Polar graphs


<< Back to FX 66-H-60 (fx66h60-il)