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NACA 66-210 (naca66210-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9


Details Polar file
Airfoil: NACA 66-210 (naca66210-il)
Reynolds number: 1,000,000
Max Cl/Cd: 82.18 at α=2.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-naca66210-il-1000000.txt
Download as CSV file: xf-naca66210-il-1000000.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 =     1.000 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.750  -0.5948   0.04229   0.03996  -0.0636   0.9702   0.0095
  -8.500  -0.5944   0.03780   0.03518  -0.0631   0.9554   0.0090
  -8.250  -0.5981   0.03189   0.02883  -0.0597   0.9415   0.0091
  -8.000  -0.6200   0.02163   0.01747  -0.0531   0.9274   0.0097
  -7.750  -0.6022   0.02020   0.01586  -0.0520   0.9206   0.0101
  -7.500  -0.5806   0.01953   0.01506  -0.0513   0.9152   0.0105
  -7.250  -0.5581   0.01882   0.01424  -0.0508   0.9099   0.0109
  -7.000  -0.5362   0.01756   0.01278  -0.0500   0.9047   0.0113
  -6.750  -0.5135   0.01629   0.01130  -0.0493   0.9001   0.0119
  -6.500  -0.4893   0.01532   0.01019  -0.0488   0.8955   0.0126
  -6.250  -0.4640   0.01494   0.00971  -0.0485   0.8910   0.0132
  -6.000  -0.4417   0.01320   0.00776  -0.0476   0.8867   0.0141
  -5.750  -0.4175   0.01246   0.00697  -0.0472   0.8824   0.0148
  -5.500  -0.3924   0.01200   0.00647  -0.0469   0.8780   0.0154
  -5.250  -0.3674   0.01152   0.00593  -0.0466   0.8741   0.0162
  -5.000  -0.3422   0.01109   0.00545  -0.0462   0.8704   0.0171
  -4.750  -0.3163   0.01077   0.00508  -0.0461   0.8666   0.0181
  -4.500  -0.2895   0.01062   0.00491  -0.0460   0.8629   0.0188
  -4.250  -0.2674   0.00973   0.00392  -0.0451   0.8592   0.0199
  -4.000  -0.2428   0.00923   0.00337  -0.0447   0.8558   0.0215
  -3.750  -0.2166   0.00896   0.00308  -0.0446   0.8522   0.0230
  -3.500  -0.1900   0.00874   0.00284  -0.0445   0.8486   0.0251
  -3.250  -0.1630   0.00861   0.00268  -0.0445   0.8455   0.0266
  -3.000  -0.1371   0.00826   0.00225  -0.0442   0.8424   0.0300
  -2.750  -0.1103   0.00803   0.00203  -0.0442   0.8394   0.0342
  -2.500  -0.0829   0.00789   0.00186  -0.0443   0.8364   0.0375
  -2.250  -0.0562   0.00764   0.00168  -0.0442   0.8335   0.0574
  -2.000  -0.0356   0.00651   0.00140  -0.0436   0.8304   0.3141
  -1.750  -0.0169   0.00517   0.00114  -0.0428   0.8273   0.6276
  -1.500   0.0088   0.00494   0.00119  -0.0424   0.8244   0.7213
  -1.250   0.0367   0.00493   0.00121  -0.0425   0.8217   0.7459
  -1.000   0.0646   0.00492   0.00123  -0.0425   0.8192   0.7620
  -0.750   0.0920   0.00496   0.00131  -0.0424   0.8168   0.7829
  -0.500   0.1199   0.00502   0.00136  -0.0424   0.8142   0.7942
  -0.250   0.1481   0.00504   0.00139  -0.0426   0.8110   0.8019
   0.000   0.1757   0.00505   0.00143  -0.0425   0.8070   0.8100
   0.250   0.2038   0.00509   0.00145  -0.0426   0.8031   0.8166
   0.500   0.2317   0.00507   0.00144  -0.0426   0.7973   0.8199
   0.750   0.2593   0.00504   0.00140  -0.0426   0.7889   0.8224
   1.000   0.2873   0.00503   0.00139  -0.0427   0.7814   0.8250
   1.250   0.3150   0.00503   0.00136  -0.0427   0.7715   0.8278
   1.500   0.3428   0.00503   0.00135  -0.0427   0.7608   0.8305
   1.750   0.3697   0.00503   0.00133  -0.0425   0.7440   0.8333
   2.000   0.3965   0.00505   0.00133  -0.0423   0.7217   0.8361
   2.250   0.4224   0.00514   0.00136  -0.0420   0.6891   0.8391
   2.500   0.4399   0.00584   0.00149  -0.0401   0.5423   0.8427
   2.750   0.4529   0.00710   0.00191  -0.0378   0.3371   0.8466
   3.000   0.4666   0.00833   0.00235  -0.0357   0.1368   0.8504
   3.250   0.4868   0.00906   0.00270  -0.0345   0.0417   0.8543
   3.500   0.5119   0.00935   0.00298  -0.0342   0.0327   0.8582
   3.750   0.5378   0.00953   0.00319  -0.0339   0.0304   0.8617
   4.000   0.5629   0.00976   0.00345  -0.0335   0.0271   0.8654
   4.250   0.5850   0.01034   0.00412  -0.0325   0.0227   0.8696
   4.500   0.6112   0.01050   0.00428  -0.0324   0.0217   0.8736
   4.750   0.6358   0.01074   0.00456  -0.0319   0.0202   0.8776
   5.000   0.6600   0.01104   0.00490  -0.0314   0.0188   0.8819
   5.250   0.6838   0.01141   0.00528  -0.0308   0.0176   0.8867
   5.500   0.7039   0.01208   0.00604  -0.0295   0.0165   0.8915
   5.750   0.7211   0.01315   0.00722  -0.0276   0.0157   0.8972
   6.250   0.7685   0.01380   0.00797  -0.0264   0.0147   0.9075
   6.500   0.7909   0.01431   0.00855  -0.0256   0.0140   0.9132
   6.750   0.8126   0.01497   0.00928  -0.0246   0.0135   0.9189
   7.000   0.8343   0.01553   0.00991  -0.0236   0.0129   0.9251
   7.250   0.8563   0.01618   0.01062  -0.0228   0.0125   0.9315
   7.500   0.8768   0.01684   0.01136  -0.0216   0.0121   0.9388
   7.750   0.8970   0.01779   0.01239  -0.0205   0.0116   0.9468
   8.000   0.9157   0.02044   0.01533  -0.0192   0.0109   0.9554
   8.250   0.9369   0.02083   0.01585  -0.0182   0.0106   0.9686
   8.500   0.9645   0.02199   0.01718  -0.0187   0.0101   0.9831
   8.750   0.9884   0.02378   0.01919  -0.0187   0.0096   1.0000
   9.000   1.0064   0.02596   0.02162  -0.0176   0.0092   1.0000
   9.250   1.0232   0.02772   0.02358  -0.0164   0.0089   1.0000
   9.500   1.0375   0.02963   0.02569  -0.0149   0.0086   1.0000
   9.750   1.0544   0.03027   0.02643  -0.0139   0.0084   1.0000
  10.000   1.0711   0.03070   0.02691  -0.0128   0.0081   1.0000
  10.250   1.0830   0.03174   0.02804  -0.0112   0.0079   1.0000
  10.500   1.0387   0.04036   0.03744  -0.0028   0.0074   1.0000
  10.750   1.0083   0.04570   0.04314   0.0034   0.0073   1.0000
  11.000   0.9765   0.05153   0.04927   0.0081   0.0072   1.0000
  11.250   0.9407   0.05767   0.05564   0.0112   0.0072   1.0000
  11.500   0.9096   0.06318   0.06131   0.0121   0.0073   1.0000
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