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NACA 66-210 (naca66210-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: NACA 66-210 (naca66210-il)
Reynolds number: 50,000
Max Cl/Cd: 22.13 at α=5.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-naca66210-il-50000.txt
Download as CSV file: xf-naca66210-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NACA 66-210                                     
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.050 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -10.000  -0.4775   0.11310   0.10594  -0.0187   1.0000   0.2837
  -9.750  -0.4757   0.11011   0.10300  -0.0183   1.0000   0.3002
  -9.500  -0.4808   0.10775   0.10072  -0.0180   1.0000   0.3170
  -9.250  -0.4698   0.10387   0.09686  -0.0169   1.0000   0.3355
  -9.000  -0.4608   0.10080   0.09381  -0.0156   1.0000   0.3575
  -8.750  -0.4520   0.09750   0.09054  -0.0141   1.0000   0.3811
  -8.500  -0.4417   0.09454   0.08760  -0.0121   1.0000   0.4114
  -8.250  -0.4321   0.09175   0.08484  -0.0101   1.0000   0.4420
  -8.000  -0.4337   0.08992   0.08305  -0.0073   1.0000   0.4750
  -7.750  -0.4216   0.08683   0.07998  -0.0055   1.0000   0.5035
  -6.250  -0.5361   0.06994   0.06403   0.0023   1.0000   0.4367
  -6.000  -0.5884   0.06776   0.06205   0.0053   1.0000   0.4330
  -5.750  -0.6297   0.05289   0.04570  -0.0181   1.0000   0.1993
  -5.500  -0.6135   0.04792   0.03982  -0.0181   1.0000   0.1575
  -5.250  -0.5974   0.04441   0.03568  -0.0167   1.0000   0.1439
  -5.000  -0.5802   0.04122   0.03203  -0.0153   1.0000   0.1381
  -4.750  -0.5595   0.03869   0.02879  -0.0138   1.0000   0.1311
  -4.500  -0.5378   0.03595   0.02567  -0.0125   1.0000   0.1266
  -4.250  -0.5141   0.03363   0.02285  -0.0112   1.0000   0.1232
  -4.000  -0.4898   0.03169   0.02046  -0.0100   1.0000   0.1229
  -3.750  -0.4661   0.03012   0.01866  -0.0089   1.0000   0.1283
  -3.500  -0.4407   0.02880   0.01698  -0.0077   1.0000   0.1328
  -3.250  -0.1396   0.02435   0.01528  -0.0368   1.0000   1.0000
  -3.000  -0.1361   0.02419   0.01488  -0.0338   1.0000   1.0000
  -2.750  -0.1324   0.02404   0.01453  -0.0308   1.0000   1.0000
  -2.500  -0.1283   0.02391   0.01417  -0.0278   1.0000   1.0000
  -2.250  -0.1240   0.02378   0.01387  -0.0249   1.0000   1.0000
  -2.000  -0.1193   0.02365   0.01359  -0.0219   1.0000   1.0000
  -1.750  -0.1145   0.02353   0.01333  -0.0190   1.0000   1.0000
  -1.500  -0.1095   0.02342   0.01308  -0.0161   1.0000   1.0000
  -1.250  -0.1044   0.02331   0.01285  -0.0132   1.0000   1.0000
  -1.000  -0.0992   0.02319   0.01262  -0.0102   1.0000   1.0000
  -0.750  -0.0940   0.02308   0.01239  -0.0073   1.0000   1.0000
  -0.500  -0.0887   0.02296   0.01218  -0.0044   1.0000   1.0000
  -0.250  -0.0833   0.02284   0.01198  -0.0015   1.0000   1.0000
   0.000  -0.0775   0.02273   0.01179   0.0014   1.0000   1.0000
   0.250  -0.0712   0.02264   0.01162   0.0042   1.0000   1.0000
   0.500  -0.0639   0.02256   0.01148   0.0068   1.0000   1.0000
   0.750  -0.0550   0.02254   0.01137   0.0091   1.0000   1.0000
   1.000  -0.0436   0.02258   0.01135   0.0110   1.0000   1.0000
   1.250  -0.0297   0.02270   0.01141   0.0124   1.0000   1.0000
   1.500  -0.0141   0.02289   0.01155   0.0135   1.0000   1.0000
   1.750   0.0028   0.02314   0.01177   0.0143   1.0000   1.0000
   2.000   0.0206   0.02344   0.01205   0.0149   1.0000   1.0000
   2.250   0.0388   0.02380   0.01241   0.0154   1.0000   1.0000
   2.500   0.0575   0.02420   0.01282   0.0158   1.0000   1.0000
   2.750   0.0762   0.02466   0.01331   0.0161   1.0000   1.0000
   3.000   0.0949   0.02516   0.01388   0.0163   1.0000   1.0000
   3.250   0.1136   0.02572   0.01450   0.0165   1.0000   1.0000
   3.500   0.1321   0.02633   0.01519   0.0166   1.0000   1.0000
   3.750   0.1504   0.02700   0.01595   0.0167   1.0000   1.0000
   4.000   0.1684   0.02773   0.01680   0.0168   1.0000   1.0000
   4.250   0.1859   0.02853   0.01772   0.0167   1.0000   1.0000
   4.500   0.2030   0.02941   0.01878   0.0167   1.0000   1.0000
   4.750   0.2194   0.03037   0.01990   0.0165   1.0000   1.0000
   5.000   0.2786   0.03284   0.02279   0.0080   0.9755   1.0000
   5.250   0.3483   0.03520   0.02574  -0.0013   0.9379   1.0000
   5.500   0.6105   0.02781   0.01610  -0.0065   0.1443   1.0000
   5.750   0.6683   0.03020   0.01839  -0.0102   0.1253   1.0000
   6.000   0.7153   0.03295   0.02140  -0.0121   0.1205   1.0000
   6.250   0.7452   0.03530   0.02413  -0.0117   0.1165   1.0000
   6.500   0.7716   0.03794   0.02691  -0.0113   0.1116   1.0000
   6.750   0.7941   0.04070   0.03021  -0.0098   0.1127   1.0000
   7.000   0.8126   0.04392   0.03405  -0.0078   0.1163   1.0000
   7.250   0.8294   0.04756   0.03812  -0.0060   0.1205   1.0000
   7.500   0.8411   0.05105   0.04228  -0.0035   0.1278   1.0000
   7.750   0.8518   0.05549   0.04715  -0.0017   0.1351   1.0000
   8.000   0.8525   0.05973   0.05205   0.0008   0.1472   1.0000
   8.250   0.8548   0.06495   0.05770   0.0022   0.1630   1.0000
   8.500   0.8457   0.07256   0.06586   0.0017   0.2064   1.0000
   8.750   0.7860   0.07848   0.07212   0.0004   0.2386   1.0000
   9.000   0.7347   0.08633   0.07991  -0.0049   0.2665   1.0000
   9.250   0.6514   0.07894   0.07282   0.0068   0.2103   1.0000
   9.500   0.6320   0.10676   0.09988  -0.0286   0.4084   1.0000
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