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NACA 0010-35 (naca001035-il) Xfoil prediction polar at RE=1,000,000 Ncrit=5


Details Polar file
Airfoil: NACA 0010-35 (naca001035-il)
Reynolds number: 1,000,000
Max Cl/Cd: 45.19 at α=6.25°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-naca001035-il-1000000-n5.txt
Download as CSV file: xf-naca001035-il-1000000-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 =     1.000 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -10.000  -0.6292   0.08015   0.07858  -0.0412   1.0000   0.0020
  -9.750  -0.6510   0.07572   0.07412  -0.0416   1.0000   0.0020
  -9.500  -0.6651   0.07035   0.06867  -0.0438   0.9985   0.0020
  -9.250  -0.6773   0.06557   0.06378  -0.0441   0.9961   0.0020
  -9.000  -0.6847   0.06062   0.05869  -0.0435   0.9932   0.0019
  -8.750  -0.6884   0.05562   0.05352  -0.0425   0.9904   0.0019
  -8.500  -0.6883   0.05025   0.04793  -0.0413   0.9884   0.0018
  -8.250  -0.6850   0.04454   0.04195  -0.0396   0.9870   0.0018
  -8.000  -0.6851   0.03915   0.03625  -0.0359   0.9830   0.0017
  -7.750  -0.6840   0.03186   0.02847  -0.0315   0.9802   0.0017
  -7.500  -0.6898   0.01867   0.01400  -0.0236   0.9780   0.0019
  -7.250  -0.6683   0.01678   0.01178  -0.0226   0.9762   0.0021
  -7.000  -0.6469   0.01559   0.01037  -0.0215   0.9737   0.0023
  -6.750  -0.6230   0.01462   0.00925  -0.0210   0.9718   0.0024
  -6.500  -0.5985   0.01376   0.00825  -0.0207   0.9702   0.0025
  -6.250  -0.5756   0.01271   0.00706  -0.0200   0.9686   0.0027
  -6.000  -0.5489   0.01234   0.00667  -0.0202   0.9673   0.0031
  -5.750  -0.5256   0.01191   0.00619  -0.0196   0.9650   0.0034
  -5.500  -0.5042   0.01138   0.00558  -0.0185   0.9618   0.0037
  -5.250  -0.4809   0.01088   0.00498  -0.0177   0.9593   0.0040
  -5.000  -0.4563   0.01046   0.00449  -0.0173   0.9571   0.0043
  -4.750  -0.4320   0.00997   0.00394  -0.0168   0.9552   0.0052
  -4.500  -0.4068   0.00969   0.00362  -0.0165   0.9532   0.0061
  -4.250  -0.3831   0.00949   0.00341  -0.0159   0.9500   0.0071
  -4.000  -0.3595   0.00916   0.00302  -0.0152   0.9469   0.0099
  -3.750  -0.3339   0.00894   0.00278  -0.0149   0.9444   0.0132
  -3.500  -0.3081   0.00868   0.00257  -0.0147   0.9423   0.0258
  -3.250  -0.2813   0.00847   0.00237  -0.0148   0.9404   0.0363
  -3.000  -0.2587   0.00820   0.00220  -0.0139   0.9368   0.0667
  -2.750  -0.2378   0.00772   0.00200  -0.0128   0.9332   0.1481
  -2.500  -0.2199   0.00697   0.00176  -0.0112   0.9297   0.2920
  -2.250  -0.1985   0.00646   0.00156  -0.0102   0.9271   0.3895
  -2.000  -0.1848   0.00575   0.00137  -0.0076   0.9223   0.5333
  -1.750  -0.1763   0.00494   0.00118  -0.0035   0.9173   0.6985
  -1.500  -0.1582   0.00457   0.00112  -0.0014   0.9140   0.7866
  -1.250  -0.1343   0.00447   0.00113  -0.0006   0.9106   0.8206
  -1.000  -0.1089   0.00442   0.00113  -0.0001   0.9070   0.8431
  -0.750  -0.0824   0.00439   0.00116   0.0001   0.9037   0.8639
  -0.500  -0.0543   0.00438   0.00116   0.0000   0.9005   0.8758
  -0.250  -0.0276   0.00438   0.00115   0.0001   0.8951   0.8826
   0.000   0.0000   0.00437   0.00114   0.0000   0.8885   0.8885
   0.250   0.0276   0.00438   0.00115  -0.0001   0.8826   0.8950
   0.500   0.0543   0.00438   0.00116   0.0000   0.8758   0.9004
   0.750   0.0823   0.00439   0.00116  -0.0001   0.8633   0.9037
   1.000   0.1089   0.00442   0.00113   0.0001   0.8427   0.9070
   1.250   0.1343   0.00447   0.00113   0.0006   0.8205   0.9107
   1.500   0.1582   0.00457   0.00113   0.0014   0.7869   0.9141
   1.750   0.1764   0.00494   0.00118   0.0035   0.6993   0.9172
   2.000   0.1851   0.00575   0.00137   0.0075   0.5347   0.9222
   2.250   0.1985   0.00646   0.00156   0.0102   0.3886   0.9271
   2.500   0.2201   0.00697   0.00176   0.0112   0.2929   0.9297
   3.000   0.2588   0.00820   0.00220   0.0139   0.0670   0.9368
   3.250   0.2815   0.00847   0.00237   0.0148   0.0363   0.9403
   3.500   0.3082   0.00868   0.00257   0.0147   0.0258   0.9423
   3.750   0.3340   0.00894   0.00278   0.0149   0.0133   0.9445
   4.000   0.3595   0.00916   0.00303   0.0151   0.0099   0.9470
   4.250   0.3832   0.00950   0.00341   0.0159   0.0071   0.9500
   4.500   0.4070   0.00969   0.00362   0.0165   0.0061   0.9531
   4.750   0.4320   0.00998   0.00395   0.0168   0.0052   0.9552
   5.000   0.4566   0.01045   0.00447   0.0172   0.0044   0.9571
   5.250   0.4809   0.01089   0.00499   0.0177   0.0040   0.9592
   5.500   0.5043   0.01138   0.00558   0.0185   0.0037   0.9618
   5.750   0.5255   0.01192   0.00620   0.0196   0.0034   0.9650
   6.000   0.5489   0.01233   0.00666   0.0202   0.0031   0.9672
   6.250   0.5753   0.01273   0.00709   0.0201   0.0028   0.9685
   6.500   0.5984   0.01375   0.00824   0.0207   0.0025   0.9701
   6.750   0.6230   0.01461   0.00923   0.0210   0.0024   0.9718
   7.000   0.6468   0.01555   0.01033   0.0215   0.0023   0.9737
   7.250   0.6681   0.01678   0.01177   0.0226   0.0021   0.9762
   7.500   0.6897   0.01866   0.01398   0.0236   0.0019   0.9780
   7.750   0.6840   0.03183   0.02844   0.0315   0.0017   0.9802
   8.000   0.6853   0.03911   0.03621   0.0359   0.0017   0.9831
   8.250   0.6852   0.04451   0.04191   0.0396   0.0018   0.9870
   8.500   0.6884   0.05026   0.04794   0.0412   0.0018   0.9884
   8.750   0.6885   0.05564   0.05354   0.0425   0.0019   0.9904
   9.000   0.6847   0.06067   0.05874   0.0435   0.0019   0.9932
   9.250   0.6774   0.06563   0.06384   0.0440   0.0020   0.9961
   9.500   0.6654   0.07039   0.06871   0.0437   0.0020   0.9985
   9.750   0.6511   0.07573   0.07413   0.0416   0.0020   1.0000
  10.000   0.6296   0.08012   0.07855   0.0412   0.0020   1.0000
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