Airfoil Tools
Search 1636 airfoils Google+
You have 0 airfoils loaded.
Your Reynold number range is 50,000 to 1,000,000. (set)

NACA 64-108 AIRFOIL (n64108-il) Xfoil prediction polar at RE=200,000 Ncrit=5


Details Polar file
Airfoil: NACA 64-108 AIRFOIL (n64108-il)
Reynolds number: 200,000
Max Cl/Cd: 43.49 at α=2.75°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-n64108-il-200000-n5.txt
Download as CSV file: xf-n64108-il-200000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NACA 64-108 AIRFOIL                             
  
 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
  ------ -------- --------- --------- -------- -------- --------
  -9.250  -0.6240   0.08622   0.08276  -0.0063   1.0000   0.0118
  -9.000  -0.6294   0.08020   0.07679  -0.0107   1.0000   0.0117
  -8.750  -0.6397   0.07303   0.06965  -0.0175   1.0000   0.0116
  -8.500  -0.6531   0.06764   0.06425  -0.0213   1.0000   0.0114
  -8.250  -0.6608   0.06218   0.05867  -0.0237   1.0000   0.0111
  -8.000  -0.6640   0.05679   0.05310  -0.0252   1.0000   0.0112
  -7.750  -0.6631   0.05143   0.04750  -0.0260   1.0000   0.0112
  -7.500  -0.6578   0.04636   0.04213  -0.0261   1.0000   0.0112
  -7.250  -0.6471   0.04208   0.03749  -0.0259   1.0000   0.0121
  -7.000  -0.6338   0.03767   0.03265  -0.0253   1.0000   0.0132
  -6.750  -0.6187   0.03348   0.02799  -0.0244   1.0000   0.0137
  -6.500  -0.6014   0.02978   0.02379  -0.0234   1.0000   0.0140
  -6.250  -0.5813   0.02696   0.02049  -0.0225   1.0000   0.0144
  -6.000  -0.5614   0.02366   0.01671  -0.0216   1.0000   0.0151
  -5.750  -0.5399   0.02175   0.01461  -0.0211   1.0000   0.0165
  -5.500  -0.5166   0.02093   0.01367  -0.0207   1.0000   0.0185
  -5.250  -0.4932   0.01944   0.01196  -0.0199   1.0000   0.0198
  -5.000  -0.4700   0.01804   0.01038  -0.0189   1.0000   0.0209
  -4.750  -0.4473   0.01690   0.00907  -0.0180   1.0000   0.0221
  -4.500  -0.4245   0.01635   0.00841  -0.0171   1.0000   0.0238
  -4.250  -0.4044   0.01505   0.00707  -0.0159   1.0000   0.0255
  -4.000  -0.3816   0.01409   0.00608  -0.0153   0.9984   0.0271
  -3.750  -0.3477   0.01331   0.00524  -0.0169   0.9902   0.0300
  -3.500  -0.3134   0.01271   0.00450  -0.0186   0.9823   0.0342
  -3.250  -0.2791   0.01213   0.00388  -0.0202   0.9742   0.0425
  -3.000  -0.2453   0.01163   0.00339  -0.0218   0.9657   0.0605
  -2.750  -0.2152   0.01039   0.00288  -0.0233   0.9557   0.2245
  -2.500  -0.1885   0.00902   0.00261  -0.0240   0.9453   0.5097
  -2.250  -0.1587   0.00870   0.00256  -0.0244   0.9355   0.5998
  -2.000  -0.1332   0.00846   0.00273  -0.0233   0.9248   0.7075
  -1.750  -0.1078   0.00842   0.00281  -0.0222   0.9132   0.7582
  -1.500  -0.0802   0.00839   0.00272  -0.0220   0.9020   0.7734
  -1.000  -0.0256   0.00833   0.00257  -0.0213   0.8800   0.7961
  -0.750   0.0011   0.00831   0.00252  -0.0208   0.8693   0.8071
  -0.500   0.0273   0.00830   0.00247  -0.0203   0.8579   0.8185
  -0.250   0.0536   0.00829   0.00245  -0.0198   0.8469   0.8300
   0.000   0.0797   0.00829   0.00243  -0.0192   0.8362   0.8419
   0.250   0.1056   0.00829   0.00243  -0.0186   0.8257   0.8538
   0.500   0.1314   0.00829   0.00245  -0.0180   0.8151   0.8656
   0.750   0.1573   0.00830   0.00248  -0.0173   0.8042   0.8779
   1.000   0.1834   0.00831   0.00251  -0.0168   0.7937   0.8906
   1.250   0.2100   0.00833   0.00256  -0.0163   0.7839   0.9037
   1.500   0.2374   0.00835   0.00263  -0.0160   0.7733   0.9172
   1.750   0.2659   0.00837   0.00270  -0.0160   0.7616   0.9309
   2.000   0.2953   0.00840   0.00273  -0.0161   0.7444   0.9448
   2.250   0.3251   0.00842   0.00268  -0.0161   0.7048   0.9586
   2.500   0.3546   0.00861   0.00257  -0.0161   0.6209   0.9734
   2.750   0.3866   0.00889   0.00260  -0.0170   0.5464   0.9902
   3.000   0.4105   0.00952   0.00272  -0.0166   0.4116   1.0000
   3.250   0.4252   0.01137   0.00328  -0.0155   0.1354   1.0000
   3.500   0.4476   0.01230   0.00382  -0.0150   0.0563   1.0000
   3.750   0.4724   0.01285   0.00432  -0.0147   0.0413   1.0000
   4.000   0.4976   0.01340   0.00495  -0.0144   0.0355   1.0000
   4.250   0.5224   0.01401   0.00564  -0.0140   0.0314   1.0000
   4.500   0.5461   0.01485   0.00652  -0.0135   0.0284   1.0000
   4.750   0.5693   0.01582   0.00754  -0.0129   0.0258   1.0000
   5.000   0.5940   0.01655   0.00840  -0.0125   0.0238   1.0000
   5.250   0.6181   0.01758   0.00952  -0.0119   0.0224   1.0000
   5.500   0.6424   0.01876   0.01081  -0.0113   0.0211   1.0000
   5.750   0.6669   0.02011   0.01229  -0.0107   0.0200   1.0000
   6.000   0.6914   0.02126   0.01355  -0.0104   0.0185   1.0000
   6.250   0.7132   0.02380   0.01626  -0.0098   0.0165   1.0000
   6.500   0.7373   0.02540   0.01821  -0.0092   0.0159   1.0000
   6.750   0.7598   0.02765   0.02087  -0.0084   0.0153   1.0000
   7.000   0.7800   0.03063   0.02432  -0.0073   0.0147   1.0000
   7.250   0.7988   0.03353   0.02766  -0.0062   0.0134   1.0000
   7.500   0.8171   0.03569   0.03013  -0.0055   0.0120   1.0000
   7.750   0.8313   0.03887   0.03367  -0.0045   0.0115   1.0000
   8.000   0.8420   0.04245   0.03761  -0.0035   0.0111   1.0000
   8.250   0.8491   0.04633   0.04182  -0.0026   0.0108   1.0000
   8.500   0.8508   0.05093   0.04675  -0.0016   0.0107   1.0000
   8.750   0.8474   0.05581   0.05193  -0.0008   0.0106   1.0000
   9.000   0.8384   0.06112   0.05749  -0.0004   0.0107   1.0000
   9.250   0.8220   0.06614   0.06269   0.0000   0.0108   1.0000
   9.500   0.7951   0.07328   0.06998  -0.0026   0.0112   1.0000
   9.750   0.7736   0.08190   0.07867  -0.0093   0.0115   1.0000
<< Back to NACA 64-108 AIRFOIL (n64108-il)

Polar data table (+)

Polar graphs


<< Back to NACA 64-108 AIRFOIL (n64108-il)