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NACA 0010-64 (naca001064-il) Xfoil prediction polar at RE=200,000 Ncrit=9


Details Polar file
Airfoil: NACA 0010-64 (naca001064-il)
Reynolds number: 200,000
Max Cl/Cd: 43.35 at α=3°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-naca001064-il-200000.txt
Download as CSV file: xf-naca001064-il-200000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NACA 0010-64                                    
  
 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
  ------ -------- --------- --------- -------- -------- --------
 -10.500  -0.5347   0.09155   0.08797  -0.0217   1.0000   0.0801
 -10.250  -0.5543   0.08443   0.08089  -0.0257   1.0000   0.0823
  -9.500  -0.7837   0.06771   0.06368  -0.0302   1.0000   0.0738
  -9.250  -0.7746   0.06489   0.06088  -0.0292   1.0000   0.0752
  -9.000  -0.8406   0.04812   0.04285  -0.0220   1.0000   0.0474
  -8.750  -0.8422   0.04422   0.03862  -0.0187   1.0000   0.0472
  -8.500  -0.8409   0.04083   0.03485  -0.0153   1.0000   0.0474
  -8.250  -0.8367   0.03757   0.03122  -0.0119   1.0000   0.0475
  -8.000  -0.8291   0.03458   0.02784  -0.0088   1.0000   0.0476
  -7.750  -0.8181   0.03238   0.02525  -0.0059   1.0000   0.0481
  -7.500  -0.8076   0.02927   0.02177  -0.0033   1.0000   0.0495
  -7.250  -0.7911   0.02740   0.01981  -0.0017   1.0000   0.0513
  -7.000  -0.7730   0.02595   0.01822   0.0000   1.0000   0.0527
  -6.750  -0.7540   0.02458   0.01667   0.0016   1.0000   0.0541
  -6.500  -0.7347   0.02353   0.01545   0.0032   1.0000   0.0564
  -6.250  -0.7146   0.02255   0.01428   0.0048   1.0000   0.0583
  -6.000  -0.6940   0.02168   0.01322   0.0064   1.0000   0.0594
  -5.750  -0.6723   0.01988   0.01139   0.0074   1.0000   0.0617
  -5.500  -0.6521   0.01908   0.01061   0.0087   1.0000   0.0645
  -5.250  -0.6316   0.01837   0.00986   0.0100   1.0000   0.0672
  -5.000  -0.6112   0.01768   0.00910   0.0115   1.0000   0.0697
  -4.750  -0.5907   0.01712   0.00848   0.0129   1.0000   0.0720
  -4.500  -0.5725   0.01616   0.00762   0.0145   1.0000   0.0770
  -4.250  -0.5526   0.01567   0.00712   0.0159   1.0000   0.0826
  -4.000  -0.5333   0.01499   0.00647   0.0174   1.0000   0.0901
  -3.750  -0.5135   0.01436   0.00593   0.0187   1.0000   0.1055
  -3.500  -0.4947   0.01341   0.00539   0.0200   1.0000   0.1656
  -3.250  -0.4783   0.01203   0.00494   0.0213   1.0000   0.3413
  -3.000  -0.4639   0.01070   0.00478   0.0235   1.0000   0.5821
  -2.750  -0.4451   0.01029   0.00492   0.0257   1.0000   0.7133
  -2.500  -0.4246   0.01026   0.00512   0.0277   1.0000   0.7840
  -2.250  -0.4044   0.01039   0.00534   0.0298   1.0000   0.8346
  -2.000  -0.3850   0.01065   0.00567   0.0323   1.0000   0.8772
  -1.750  -0.3642   0.01102   0.00605   0.0346   1.0000   0.9129
  -1.500  -0.3349   0.01144   0.00643   0.0349   1.0000   0.9380
  -1.250  -0.2966   0.01183   0.00674   0.0329   0.9997   0.9551
  -1.000  -0.2325   0.01248   0.00731   0.0261   1.0000   0.9705
  -0.750  -0.1528   0.01300   0.00775   0.0160   1.0000   0.9804
  -0.500  -0.0926   0.01318   0.00789   0.0091   1.0000   0.9881
  -0.250  -0.0460   0.01326   0.00794   0.0045   1.0000   0.9947
   0.000   0.0000   0.01330   0.00798   0.0000   1.0000   1.0000
   0.250   0.0460   0.01326   0.00794  -0.0045   0.9947   1.0000
   0.500   0.0926   0.01318   0.00789  -0.0091   0.9881   1.0000
   0.750   0.1527   0.01300   0.00774  -0.0159   0.9804   1.0000
   1.000   0.2324   0.01248   0.00731  -0.0261   0.9705   1.0000
   1.250   0.2966   0.01183   0.00674  -0.0329   0.9552   0.9997
   1.500   0.3348   0.01144   0.00643  -0.0349   0.9379   1.0000
   1.750   0.3640   0.01102   0.00605  -0.0346   0.9131   1.0000
   2.000   0.3849   0.01065   0.00566  -0.0323   0.8771   1.0000
   2.250   0.4043   0.01038   0.00534  -0.0298   0.8348   1.0000
   2.500   0.4244   0.01026   0.00511  -0.0277   0.7839   1.0000
   2.750   0.4450   0.01029   0.00492  -0.0257   0.7135   1.0000
   3.000   0.4638   0.01070   0.00478  -0.0234   0.5822   1.0000
   3.250   0.4782   0.01203   0.00494  -0.0213   0.3419   1.0000
   3.500   0.4946   0.01341   0.00538  -0.0200   0.1656   1.0000
   3.750   0.5134   0.01435   0.00593  -0.0187   0.1056   1.0000
   4.000   0.5333   0.01498   0.00647  -0.0174   0.0901   1.0000
   4.250   0.5525   0.01567   0.00712  -0.0159   0.0826   1.0000
   4.500   0.5724   0.01616   0.00762  -0.0145   0.0770   1.0000
   4.750   0.5906   0.01712   0.00848  -0.0129   0.0720   1.0000
   5.000   0.6111   0.01768   0.00910  -0.0114   0.0697   1.0000
   5.250   0.6315   0.01837   0.00986  -0.0100   0.0672   1.0000
   5.500   0.6520   0.01908   0.01061  -0.0086   0.0645   1.0000
   5.750   0.6722   0.01988   0.01138  -0.0073   0.0617   1.0000
   6.000   0.6939   0.02168   0.01322  -0.0063   0.0594   1.0000
   6.250   0.7145   0.02255   0.01428  -0.0048   0.0583   1.0000
   6.500   0.7346   0.02353   0.01545  -0.0032   0.0564   1.0000
   6.750   0.7539   0.02457   0.01667  -0.0016   0.0541   1.0000
   7.000   0.7729   0.02595   0.01822   0.0000   0.0527   1.0000
   7.250   0.7910   0.02739   0.01980   0.0017   0.0513   1.0000
   7.500   0.8075   0.02926   0.02177   0.0033   0.0495   1.0000
   7.750   0.8181   0.03237   0.02525   0.0059   0.0481   1.0000
   8.000   0.8290   0.03457   0.02783   0.0088   0.0476   1.0000
   8.250   0.8366   0.03757   0.03121   0.0119   0.0475   1.0000
   8.500   0.8409   0.04083   0.03485   0.0153   0.0474   1.0000
   8.750   0.8422   0.04422   0.03863   0.0187   0.0472   1.0000
   9.000   0.8406   0.04813   0.04286   0.0220   0.0474   1.0000
   9.250   0.6752   0.05704   0.05335   0.0356   0.0817   1.0000
   9.500   0.6452   0.06151   0.05790   0.0373   0.0817   1.0000
   9.750   0.6162   0.06722   0.06367   0.0363   0.0823   1.0000
  10.000   0.5869   0.07459   0.07108   0.0326   0.0829   1.0000
  10.250   0.5542   0.08429   0.08075   0.0257   0.0824   1.0000
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