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NACA 0010-35 (naca001035-il) Xfoil prediction polar at RE=200,000 Ncrit=5


Details Polar file
Airfoil: NACA 0010-35 (naca001035-il)
Reynolds number: 200,000
Max Cl/Cd: 32.22 at α=2.5°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-naca001035-il-200000-n5.txt
Download as CSV file: xf-naca001035-il-200000-n5.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 =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -10.000  -0.5941   0.08976   0.08623  -0.0368   1.0000   0.0104
  -9.750  -0.6065   0.08386   0.08035  -0.0408   1.0000   0.0102
  -9.500  -0.6227   0.07917   0.07564  -0.0421   1.0000   0.0101
  -9.250  -0.6418   0.07548   0.07191  -0.0410   1.0000   0.0100
  -9.000  -0.6630   0.07263   0.06903  -0.0377   1.0000   0.0099
  -8.750  -0.6859   0.07045   0.06682  -0.0325   1.0000   0.0098
  -8.500  -0.7045   0.06774   0.06404  -0.0277   1.0000   0.0096
  -8.250  -0.7202   0.06475   0.06095  -0.0230   1.0000   0.0093
  -8.000  -0.7336   0.06154   0.05761  -0.0182   1.0000   0.0090
  -7.750  -0.7450   0.05803   0.05392  -0.0133   1.0000   0.0087
  -7.500  -0.7542   0.05407   0.04974  -0.0081   1.0000   0.0082
  -7.250  -0.7586   0.05084   0.04629  -0.0035   1.0000   0.0082
  -7.000  -0.7579   0.04833   0.04358   0.0002   0.9999   0.0087
  -6.750  -0.7434   0.04482   0.03975   0.0013   0.9974   0.0094
  -6.500  -0.7271   0.04103   0.03557   0.0028   0.9951   0.0099
  -6.250  -0.7123   0.03713   0.03125   0.0050   0.9924   0.0098
  -6.000  -0.6941   0.03332   0.02695   0.0069   0.9900   0.0095
  -5.750  -0.6721   0.02989   0.02301   0.0084   0.9881   0.0094
  -5.500  -0.6469   0.02697   0.01960   0.0093   0.9869   0.0095
  -5.250  -0.6225   0.02465   0.01689   0.0104   0.9852   0.0097
  -5.000  -0.5958   0.02271   0.01463   0.0110   0.9836   0.0103
  -4.750  -0.5679   0.02159   0.01325   0.0110   0.9819   0.0118
  -4.500  -0.5406   0.02012   0.01160   0.0111   0.9805   0.0129
  -4.250  -0.5164   0.01845   0.00982   0.0117   0.9792   0.0143
  -4.000  -0.4894   0.01758   0.00887   0.0116   0.9776   0.0169
  -3.750  -0.4673   0.01699   0.00817   0.0126   0.9744   0.0209
  -3.500  -0.4438   0.01623   0.00728   0.0132   0.9716   0.0263
  -3.250  -0.4169   0.01566   0.00665   0.0131   0.9694   0.0414
  -3.000  -0.3929   0.01465   0.00612   0.0132   0.9675   0.1403
  -2.750  -0.3899   0.01259   0.00558   0.0172   0.9630   0.4797
  -2.500  -0.3723   0.01155   0.00597   0.0203   0.9617   0.7958
  -2.250  -0.3305   0.01236   0.00684   0.0188   0.9625   0.8867
  -2.000  -0.2905   0.01319   0.00755   0.0172   0.9626   0.9151
  -1.750  -0.2541   0.01360   0.00780   0.0156   0.9614   0.9271
  -1.500  -0.2136   0.01375   0.00784   0.0126   0.9604   0.9295
  -1.250  -0.1733   0.01390   0.00788   0.0096   0.9594   0.9329
  -1.000  -0.1397   0.01397   0.00786   0.0080   0.9575   0.9385
  -0.750  -0.1037   0.01400   0.00783   0.0058   0.9559   0.9416
  -0.500  -0.0645   0.01402   0.00781   0.0028   0.9549   0.9431
  -0.250  -0.0295   0.01404   0.00780   0.0008   0.9526   0.9453
   0.000   0.0000   0.01406   0.00782   0.0000   0.9486   0.9486
   0.250   0.0295   0.01404   0.00780  -0.0008   0.9453   0.9526
   0.500   0.0645   0.01402   0.00781  -0.0028   0.9431   0.9549
   0.750   0.1037   0.01400   0.00783  -0.0058   0.9416   0.9559
   1.000   0.1397   0.01397   0.00786  -0.0080   0.9386   0.9575
   1.250   0.1733   0.01390   0.00788  -0.0096   0.9329   0.9594
   1.500   0.2136   0.01375   0.00784  -0.0126   0.9295   0.9604
   1.750   0.2542   0.01359   0.00780  -0.0156   0.9271   0.9614
   2.000   0.2913   0.01316   0.00752  -0.0174   0.9148   0.9624
   2.250   0.3306   0.01236   0.00683  -0.0188   0.8867   0.9625
   2.500   0.3721   0.01155   0.00597  -0.0202   0.7965   0.9617
   2.750   0.3898   0.01259   0.00558  -0.0172   0.4799   0.9630
   3.000   0.3929   0.01465   0.00612  -0.0132   0.1404   0.9674
   3.250   0.4169   0.01566   0.00665  -0.0131   0.0415   0.9694
   3.500   0.4437   0.01623   0.00728  -0.0132   0.0262   0.9716
   3.750   0.4673   0.01699   0.00817  -0.0126   0.0209   0.9744
   4.000   0.4895   0.01758   0.00887  -0.0116   0.0169   0.9777
   4.250   0.5165   0.01845   0.00982  -0.0118   0.0143   0.9792
   4.500   0.5405   0.02013   0.01160  -0.0111   0.0129   0.9806
   4.750   0.5678   0.02163   0.01328  -0.0109   0.0119   0.9819
   5.000   0.5957   0.02271   0.01463  -0.0109   0.0103   0.9836
   5.250   0.6224   0.02465   0.01689  -0.0104   0.0097   0.9852
   5.500   0.6469   0.02696   0.01960  -0.0093   0.0095   0.9869
   5.750   0.6721   0.02987   0.02299  -0.0084   0.0094   0.9881
   6.000   0.6941   0.03332   0.02695  -0.0069   0.0095   0.9900
   6.250   0.7123   0.03712   0.03124  -0.0050   0.0098   0.9924
   6.500   0.7271   0.04103   0.03558  -0.0027   0.0099   0.9951
   6.750   0.7434   0.04480   0.03974  -0.0014   0.0094   0.9974
   7.000   0.7581   0.04824   0.04349  -0.0003   0.0086   0.9999
   7.250   0.7586   0.05081   0.04625   0.0035   0.0082   1.0000
   7.500   0.7544   0.05399   0.04965   0.0081   0.0082   1.0000
   7.750   0.7450   0.05800   0.05390   0.0133   0.0086   1.0000
   8.000   0.7337   0.06153   0.05760   0.0182   0.0090   1.0000
   8.250   0.7203   0.06475   0.06094   0.0230   0.0093   1.0000
   8.500   0.7046   0.06773   0.06403   0.0277   0.0096   1.0000
   8.750   0.6860   0.07045   0.06682   0.0325   0.0098   1.0000
   9.000   0.6631   0.07264   0.06904   0.0377   0.0099   1.0000
   9.250   0.6420   0.07548   0.07192   0.0410   0.0100   1.0000
   9.500   0.6231   0.07916   0.07563   0.0421   0.0101   1.0000
   9.750   0.6068   0.08386   0.08035   0.0407   0.0102   1.0000
  10.000   0.5946   0.08976   0.08623   0.0368   0.0104   1.0000
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