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A18 (original) (a18-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: A18 (original) (a18-il)
Reynolds number: 50,000
Max Cl/Cd: 43.99 at α=5.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-a18-il-50000.txt
Download as CSV file: xf-a18-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: A18 (original)                                  
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.050 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.500  -0.3172   0.10220   0.09553  -0.0249   1.0000   0.1971
  -9.250  -0.3132   0.09842   0.09180  -0.0242   1.0000   0.2036
  -9.000  -0.3191   0.09646   0.08992  -0.0239   1.0000   0.2120
  -8.750  -0.3121   0.09236   0.08585  -0.0228   1.0000   0.2195
  -8.500  -0.3236   0.09086   0.08447  -0.0222   1.0000   0.2269
  -8.250  -0.3159   0.08659   0.08023  -0.0210   1.0000   0.2338
  -8.000  -0.3296   0.08519   0.07897  -0.0195   1.0000   0.2411
  -7.750  -0.3235   0.08118   0.07500  -0.0180   1.0000   0.2487
  -7.500  -0.4116   0.09141   0.08489  -0.0209   1.0000   0.2429
  -7.250  -0.4000   0.08715   0.08064  -0.0174   1.0000   0.2543
  -7.000  -0.3951   0.08382   0.07737  -0.0164   1.0000   0.2636
  -6.750  -0.3924   0.08102   0.07462  -0.0161   1.0000   0.2768
  -6.500  -0.3883   0.07832   0.07198  -0.0151   1.0000   0.2948
  -6.250  -0.3867   0.07624   0.06996  -0.0176   1.0000   0.3152
  -6.000  -0.3798   0.07281   0.06654  -0.0142   1.0000   0.3319
  -5.750  -0.3729   0.06986   0.06362  -0.0122   1.0000   0.3503
  -5.500  -0.3673   0.06718   0.06098  -0.0115   1.0000   0.3756
  -5.250  -0.3621   0.06447   0.05832  -0.0082   1.0000   0.4055
  -5.000  -0.3568   0.06180   0.05568  -0.0050   1.0000   0.4364
  -4.500  -0.1719   0.03986   0.03168  -0.0687   1.0000   0.2095
  -4.250  -0.1231   0.03473   0.02556  -0.0762   1.0000   0.1985
  -4.000  -0.0973   0.03293   0.02364  -0.0767   1.0000   0.2065
  -3.750  -0.0621   0.03061   0.02073  -0.0792   1.0000   0.2117
  -3.500  -0.0300   0.02871   0.01843  -0.0808   1.0000   0.2161
  -3.250  -0.0006   0.02738   0.01681  -0.0815   1.0000   0.2250
  -3.000   0.0284   0.02640   0.01554  -0.0822   1.0000   0.2421
  -2.750   0.0576   0.02537   0.01439  -0.0828   1.0000   0.2585
  -2.500   0.0859   0.02451   0.01352  -0.0831   1.0000   0.2808
  -2.250   0.1133   0.02402   0.01310  -0.0832   1.0000   0.3212
  -2.000   0.1426   0.02342   0.01252  -0.0835   1.0000   0.3867
  -1.750   0.1719   0.02244   0.01205  -0.0838   1.0000   0.4909
  -1.500   0.1950   0.02068   0.01163  -0.0819   1.0000   0.7351
  -1.250   0.2070   0.02060   0.01129  -0.0799   1.0000   1.0000
  -1.000   0.2317   0.02115   0.01146  -0.0803   1.0000   1.0000
  -0.750   0.2549   0.02175   0.01179  -0.0804   1.0000   1.0000
  -0.500   0.2771   0.02241   0.01226  -0.0806   1.0000   1.0000
  -0.250   0.2988   0.02313   0.01285  -0.0807   1.0000   1.0000
   0.000   0.3197   0.02393   0.01353  -0.0808   1.0000   1.0000
   0.250   0.3401   0.02481   0.01432  -0.0810   1.0000   1.0000
   0.500   0.3598   0.02576   0.01521  -0.0812   1.0000   1.0000
   0.750   0.3789   0.02679   0.01620  -0.0814   1.0000   1.0000
   1.000   0.3975   0.02790   0.01729  -0.0817   1.0000   1.0000
   1.250   0.4312   0.02909   0.01847  -0.0848   0.9935   1.0000
   1.500   0.4769   0.03021   0.01962  -0.0900   0.9788   1.0000
   1.750   0.5207   0.03129   0.02075  -0.0946   0.9639   1.0000
   2.000   0.5628   0.03231   0.02188  -0.0987   0.9489   1.0000
   2.250   0.6029   0.03331   0.02298  -0.1023   0.9341   1.0000
   2.500   0.6428   0.03424   0.02405  -0.1057   0.9188   1.0000
   2.750   0.6944   0.03476   0.02479  -0.1103   0.9005   1.0000
   3.000   0.7372   0.03487   0.02510  -0.1126   0.8753   1.0000
   3.250   0.7899   0.03421   0.02471  -0.1152   0.8479   1.0000
   3.500   0.8410   0.03343   0.02427  -0.1173   0.8249   1.0000
   3.750   0.8877   0.03269   0.02388  -0.1186   0.8038   1.0000
   4.000   0.9309   0.03172   0.02330  -0.1187   0.7811   1.0000
   4.250   0.9806   0.02988   0.02186  -0.1184   0.7560   1.0000
   4.500   1.0155   0.02848   0.02080  -0.1157   0.7213   1.0000
   4.750   1.0535   0.02593   0.01850  -0.1112   0.6696   1.0000
   5.000   1.0739   0.02466   0.01720  -0.1053   0.5965   1.0000
   5.250   1.0853   0.02467   0.01680  -0.0991   0.4889   1.0000
   5.500   1.0873   0.02616   0.01728  -0.0930   0.3576   1.0000
   5.750   1.0872   0.02915   0.01900  -0.0883   0.2427   1.0000
   6.000   1.1021   0.03323   0.02227  -0.0858   0.1642   1.0000
   6.250   1.1282   0.03621   0.02509  -0.0846   0.1315   1.0000
   6.500   1.1546   0.03886   0.02764  -0.0840   0.1140   1.0000
   6.750   1.1837   0.04220   0.03137  -0.0832   0.1058   1.0000
   7.000   1.2087   0.04561   0.03483  -0.0826   0.0990   1.0000
   7.250   1.2257   0.04897   0.03887  -0.0805   0.0943   1.0000
   7.500   1.2416   0.05288   0.04329  -0.0786   0.0921   1.0000
   7.750   1.2520   0.05732   0.04829  -0.0762   0.0919   1.0000
   8.000   1.2563   0.06213   0.05368  -0.0736   0.0928   1.0000
   8.250   1.2553   0.06715   0.05922  -0.0708   0.0943   1.0000
   8.500   1.2498   0.07221   0.06471  -0.0682   0.0957   1.0000
   8.750   1.2419   0.07724   0.07007  -0.0657   0.0970   1.0000
   9.000   1.2332   0.08232   0.07540  -0.0637   0.0982   1.0000
   9.250   1.2265   0.08758   0.08087  -0.0621   0.0992   1.0000
   9.500   1.1644   0.09253   0.08624  -0.0592   0.1038   1.0000
   9.750   1.1246   0.09954   0.09341  -0.0610   0.1063   1.0000
  10.000   1.0957   0.10759   0.10152  -0.0649   0.1090   1.0000
  10.250   0.9240   0.10792   0.10207  -0.0523   0.1101   1.0000
  10.500   0.9119   0.11446   0.10861  -0.0540   0.1121   1.0000
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