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NACA 0010-35 (naca001035-il) Xfoil prediction polar at RE=200,000 Ncrit=9


Details Polar file
Airfoil: NACA 0010-35 (naca001035-il)
Reynolds number: 200,000
Max Cl/Cd: 37.42 at α=2.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-naca001035-il-200000.txt
Download as CSV file: xf-naca001035-il-200000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NACA 0010-35                                    
  
 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
  ------ -------- --------- --------- -------- -------- --------
 -10.250  -0.4994   0.08976   0.08642  -0.0396   1.0000   0.0419
 -10.000  -0.5131   0.08395   0.08064  -0.0424   1.0000   0.0419
  -9.750  -0.5294   0.07834   0.07502  -0.0445   1.0000   0.0417
  -9.500  -0.5479   0.07357   0.07024  -0.0451   1.0000   0.0414
  -9.250  -0.5691   0.06943   0.06608  -0.0443   1.0000   0.0411
  -9.000  -0.5919   0.06597   0.06259  -0.0419   1.0000   0.0408
  -8.750  -0.6162   0.06318   0.05977  -0.0379   1.0000   0.0406
  -8.500  -0.6415   0.06101   0.05757  -0.0323   1.0000   0.0405
  -8.250  -0.6631   0.05856   0.05505  -0.0271   1.0000   0.0407
  -8.000  -0.6818   0.05608   0.05243  -0.0219   1.0000   0.0414
  -7.750  -0.7016   0.05500   0.05106  -0.0154   1.0000   0.0426
  -7.500  -0.7136   0.05383   0.04961  -0.0097   1.0000   0.0430
  -7.250  -0.7218   0.05163   0.04717  -0.0048   1.0000   0.0432
  -7.000  -0.7702   0.05688   0.05180   0.0020   1.0000   0.0436
  -6.750  -0.7696   0.05052   0.04551   0.0039   1.0000   0.0451
  -6.500  -0.7637   0.04768   0.04264   0.0065   1.0000   0.0467
  -6.250  -0.7577   0.04529   0.04011   0.0096   1.0000   0.0490
  -6.000  -0.7505   0.04727   0.04140   0.0153   1.0000   0.0558
  -5.750  -0.7474   0.04043   0.03454   0.0174   1.0000   0.0583
  -5.500  -0.7351   0.03798   0.03205   0.0194   1.0000   0.0616
  -5.250  -0.7267   0.03619   0.02983   0.0230   1.0000   0.0713
  -5.000  -0.7112   0.03385   0.02742   0.0247   1.0000   0.0764
  -4.750  -0.6787   0.02803   0.02057   0.0291   1.0000   0.0378
  -4.500  -0.6543   0.02535   0.01744   0.0309   1.0000   0.0355
  -4.250  -0.6293   0.02344   0.01522   0.0322   1.0000   0.0356
  -4.000  -0.6018   0.02152   0.01300   0.0331   1.0000   0.0340
  -3.750  -0.5773   0.02010   0.01147   0.0341   1.0000   0.0349
  -3.500  -0.5552   0.01898   0.01030   0.0354   1.0000   0.0369
  -3.250  -0.5345   0.01832   0.00957   0.0367   1.0000   0.0418
  -3.000  -0.5159   0.01719   0.00838   0.0386   1.0000   0.0455
  -2.750  -0.4972   0.01643   0.00761   0.0402   1.0000   0.0520
  -2.500  -0.4776   0.01581   0.00695   0.0417   1.0000   0.0672
  -2.250  -0.4725   0.01265   0.00605   0.0452   1.0000   0.5224
  -2.000  -0.2298   0.01702   0.01144   0.0091   1.0000   0.9738
  -1.750  -0.1802   0.01712   0.01140   0.0040   1.0000   0.9830
  -1.500  -0.1373   0.01699   0.01114   0.0000   1.0000   0.9872
  -1.250  -0.0977   0.01692   0.01099  -0.0034   1.0000   0.9928
  -1.000  -0.0552   0.01681   0.01083  -0.0073   1.0000   0.9974
  -0.750  -0.0266   0.01673   0.01071  -0.0085   1.0000   1.0000
  -0.500  -0.0177   0.01667   0.01064  -0.0057   1.0000   1.0000
  -0.250  -0.0089   0.01664   0.01060  -0.0029   1.0000   1.0000
   0.000   0.0000   0.01663   0.01058   0.0000   1.0000   1.0000
   0.250   0.0089   0.01664   0.01060   0.0029   1.0000   1.0000
   0.500   0.0177   0.01667   0.01064   0.0057   1.0000   1.0000
   0.750   0.0266   0.01672   0.01070   0.0085   1.0000   1.0000
   1.000   0.0552   0.01680   0.01082   0.0073   0.9974   1.0000
   1.250   0.0977   0.01691   0.01099   0.0034   0.9928   1.0000
   1.500   0.1372   0.01698   0.01114   0.0000   0.9872   1.0000
   1.750   0.1801   0.01712   0.01140  -0.0040   0.9830   1.0000
   2.000   0.2305   0.01701   0.01143  -0.0092   0.9737   1.0000
   2.250   0.4726   0.01263   0.00605  -0.0452   0.5262   1.0000
   2.500   0.4776   0.01580   0.00695  -0.0417   0.0673   1.0000
   2.750   0.4972   0.01643   0.00760  -0.0402   0.0521   1.0000
   3.000   0.5159   0.01719   0.00837  -0.0386   0.0455   1.0000
   3.250   0.5345   0.01831   0.00956  -0.0367   0.0418   1.0000
   3.500   0.5552   0.01898   0.01029  -0.0354   0.0369   1.0000
   3.750   0.5773   0.02010   0.01147  -0.0341   0.0348   1.0000
   4.000   0.6017   0.02151   0.01299  -0.0331   0.0340   1.0000
   4.250   0.6292   0.02344   0.01521  -0.0322   0.0356   1.0000
   4.500   0.6543   0.02536   0.01746  -0.0309   0.0356   1.0000
   4.750   0.6786   0.02802   0.02056  -0.0291   0.0378   1.0000
   5.000   0.6927   0.02149   0.01556  -0.0229   0.0754   1.0000
   5.250   0.7067   0.02421   0.01832  -0.0213   0.0710   1.0000
   5.500   0.7120   0.02608   0.02064  -0.0173   0.0617   1.0000
   5.750   0.7224   0.02867   0.02327  -0.0151   0.0583   1.0000
   7.000   0.7702   0.05685   0.05177  -0.0020   0.0436   1.0000
   7.250   0.7217   0.05153   0.04707   0.0048   0.0432   1.0000
   7.500   0.7135   0.05375   0.04953   0.0097   0.0430   1.0000
   7.750   0.7011   0.05485   0.05091   0.0155   0.0425   1.0000
   8.000   0.6817   0.05598   0.05233   0.0219   0.0414   1.0000
   8.250   0.6630   0.05843   0.05492   0.0270   0.0407   1.0000
   8.500   0.6413   0.06088   0.05744   0.0323   0.0405   1.0000
   8.750   0.6161   0.06306   0.05964   0.0379   0.0406   1.0000
   9.000   0.5918   0.06584   0.06246   0.0419   0.0408   1.0000
   9.250   0.5688   0.06931   0.06596   0.0443   0.0411   1.0000
   9.500   0.5482   0.07339   0.07007   0.0452   0.0414   1.0000
   9.750   0.5292   0.07822   0.07490   0.0445   0.0417   1.0000
  10.000   0.5128   0.08381   0.08050   0.0424   0.0419   1.0000
  10.250   0.4989   0.08965   0.08631   0.0395   0.0419   1.0000
  10.500   0.4871   0.09498   0.09160   0.0372   0.0415   1.0000
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