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NACA 0010-34 (naca001034-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9


Details Polar file
Airfoil: NACA 0010-34 (naca001034-il)
Reynolds number: 1,000,000
Max Cl/Cd: 49.65 at α=2.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-naca001034-il-1000000.txt
Download as CSV file: xf-naca001034-il-1000000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NACA 0010-34                                    
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     1.000 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -10.750  -0.6799   0.08213   0.08059  -0.0229   1.0000   0.0068
 -10.500  -0.6942   0.07141   0.06976  -0.0317   1.0000   0.0066
 -10.250  -0.7066   0.06573   0.06398  -0.0352   1.0000   0.0066
 -10.000  -0.7233   0.06097   0.05910  -0.0366   1.0000   0.0066
  -9.750  -0.7422   0.05689   0.05489  -0.0357   1.0000   0.0066
  -9.500  -0.7668   0.05299   0.05085  -0.0321   1.0000   0.0066
  -9.250  -0.7845   0.04890   0.04657  -0.0287   1.0000   0.0067
  -9.000  -0.7937   0.04574   0.04324  -0.0254   1.0000   0.0069
  -8.750  -0.7954   0.04372   0.04110  -0.0223   1.0000   0.0071
  -8.500  -0.7935   0.04223   0.03952  -0.0194   1.0000   0.0073
  -8.250  -0.7910   0.04082   0.03801  -0.0162   1.0000   0.0076
  -8.000  -0.7889   0.03917   0.03624  -0.0126   1.0000   0.0081
  -7.250  -0.7826   0.02276   0.01828   0.0011   0.9976   0.0068
  -7.000  -0.7569   0.02047   0.01573   0.0008   0.9963   0.0071
  -6.750  -0.7261   0.02046   0.01574  -0.0009   0.9951   0.0078
  -6.500  -0.7008   0.01648   0.01131   0.0000   0.9944   0.0072
  -6.250  -0.6741   0.01500   0.00967   0.0000   0.9926   0.0074
  -6.000  -0.6455   0.01410   0.00870  -0.0005   0.9906   0.0079
  -5.750  -0.6154   0.01339   0.00793  -0.0014   0.9887   0.0087
  -5.500  -0.5849   0.01266   0.00714  -0.0023   0.9870   0.0095
  -5.250  -0.5535   0.01201   0.00643  -0.0033   0.9856   0.0100
  -5.000  -0.5209   0.01152   0.00587  -0.0047   0.9844   0.0105
  -4.750  -0.4913   0.01031   0.00452  -0.0055   0.9832   0.0122
  -4.500  -0.4592   0.00988   0.00406  -0.0067   0.9816   0.0142
  -4.250  -0.4312   0.00954   0.00368  -0.0069   0.9778   0.0159
  -4.000  -0.4003   0.00925   0.00337  -0.0078   0.9749   0.0172
  -3.750  -0.3697   0.00878   0.00284  -0.0085   0.9722   0.0255
  -3.500  -0.3415   0.00810   0.00244  -0.0090   0.9694   0.0966
  -3.250  -0.3236   0.00706   0.00206  -0.0075   0.9624   0.2638
  -3.000  -0.3014   0.00630   0.00180  -0.0068   0.9572   0.3956
  -2.750  -0.2802   0.00570   0.00161  -0.0056   0.9508   0.5059
  -2.500  -0.2587   0.00521   0.00144  -0.0044   0.9442   0.6022
  -2.250  -0.2368   0.00484   0.00135  -0.0032   0.9372   0.6826
  -2.000  -0.2131   0.00462   0.00128  -0.0022   0.9300   0.7359
  -1.750  -0.1878   0.00450   0.00123  -0.0016   0.9226   0.7656
  -1.500  -0.1619   0.00441   0.00118  -0.0011   0.9156   0.7916
  -1.250  -0.1351   0.00436   0.00114  -0.0009   0.9080   0.8053
  -1.000  -0.1081   0.00432   0.00110  -0.0007   0.9008   0.8191
  -0.750  -0.0814   0.00428   0.00108  -0.0005   0.8922   0.8329
  -0.500  -0.0543   0.00426   0.00106  -0.0003   0.8842   0.8455
  -0.250  -0.0272   0.00424   0.00105  -0.0001   0.8752   0.8561
   0.000   0.0000   0.00423   0.00105   0.0000   0.8659   0.8658
   0.250   0.0273   0.00424   0.00105   0.0001   0.8562   0.8752
   0.500   0.0543   0.00426   0.00106   0.0003   0.8454   0.8842
   0.750   0.0814   0.00428   0.00108   0.0005   0.8330   0.8922
   1.000   0.1082   0.00432   0.00110   0.0007   0.8191   0.9008
   1.250   0.1352   0.00436   0.00114   0.0009   0.8053   0.9079
   1.500   0.1620   0.00441   0.00118   0.0011   0.7916   0.9156
   1.750   0.1880   0.00449   0.00123   0.0016   0.7669   0.9226
   2.000   0.2132   0.00462   0.00128   0.0022   0.7358   0.9300
   2.250   0.2370   0.00483   0.00135   0.0031   0.6845   0.9371
   2.500   0.2587   0.00521   0.00144   0.0044   0.6014   0.9442
   2.750   0.2802   0.00571   0.00161   0.0056   0.5054   0.9508
   3.000   0.3014   0.00630   0.00180   0.0067   0.3958   0.9572
   3.250   0.3236   0.00707   0.00206   0.0075   0.2636   0.9624
   3.500   0.3415   0.00810   0.00244   0.0090   0.0967   0.9694
   3.750   0.3697   0.00878   0.00284   0.0085   0.0256   0.9722
   4.000   0.4002   0.00926   0.00337   0.0078   0.0172   0.9749
   4.250   0.4312   0.00953   0.00368   0.0069   0.0159   0.9777
   4.500   0.4593   0.00987   0.00405   0.0067   0.0142   0.9816
   4.750   0.4912   0.01032   0.00453   0.0055   0.0123   0.9832
   5.000   0.5207   0.01153   0.00588   0.0047   0.0105   0.9844
   5.250   0.5534   0.01203   0.00645   0.0034   0.0101   0.9856
   5.500   0.5848   0.01268   0.00715   0.0023   0.0095   0.9870
   5.750   0.6154   0.01338   0.00792   0.0013   0.0087   0.9887
   6.000   0.6453   0.01411   0.00871   0.0006   0.0080   0.9906
   6.250   0.6739   0.01502   0.00969   0.0001   0.0075   0.9926
   6.500   0.7007   0.01649   0.01132   0.0000   0.0073   0.9944
   6.750   0.7277   0.01945   0.01460   0.0002   0.0075   0.9951
   7.000   0.7571   0.02041   0.01567  -0.0009   0.0070   0.9963
   7.250   0.7828   0.02278   0.01830  -0.0011   0.0068   0.9977
   9.250   0.7204   0.03934   0.03723   0.0347   0.0071   1.0000
   9.500   0.7070   0.04120   0.03916   0.0385   0.0070   1.0000
   9.750   0.6830   0.04504   0.04314   0.0415   0.0070   1.0000
  10.000   0.6560   0.04972   0.04796   0.0425   0.0070   1.0000
  10.250   0.6119   0.05880   0.05721   0.0399   0.0073   1.0000
  10.500   0.5554   0.07669   0.07521   0.0284   0.0086   1.0000
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