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NACA 66-209 (naca66209-il) Xfoil prediction polar at RE=200,000 Ncrit=5


Details Polar file
Airfoil: NACA 66-209 (naca66209-il)
Reynolds number: 200,000
Max Cl/Cd: 47.11 at α=3°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-naca66209-il-200000-n5.txt
Download as CSV file: xf-naca66209-il-200000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NACA 66-209                                     
  
 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
  ------ -------- --------- --------- -------- -------- --------
  -9.500  -0.5149   0.08660   0.08302  -0.0368   1.0000   0.0160
  -9.250  -0.5197   0.08142   0.07789  -0.0402   1.0000   0.0160
  -9.000  -0.5271   0.07621   0.07275  -0.0443   1.0000   0.0157
  -8.750  -0.5391   0.07184   0.06839  -0.0465   1.0000   0.0153
  -8.500  -0.5553   0.06839   0.06494  -0.0461   1.0000   0.0153
  -8.250  -0.5696   0.06492   0.06144  -0.0446   1.0000   0.0151
  -8.000  -0.5836   0.06224   0.05873  -0.0418   1.0000   0.0148
  -7.750  -0.5813   0.05713   0.05343  -0.0434   0.9939   0.0145
  -7.500  -0.5686   0.05116   0.04716  -0.0462   0.9868   0.0141
  -7.250  -0.5540   0.04563   0.04124  -0.0479   0.9800   0.0139
  -7.000  -0.5352   0.04025   0.03542  -0.0493   0.9754   0.0140
  -6.750  -0.5166   0.03593   0.03068  -0.0495   0.9692   0.0141
  -6.500  -0.4929   0.03180   0.02603  -0.0500   0.9653   0.0149
  -6.250  -0.4699   0.02898   0.02273  -0.0495   0.9596   0.0165
  -6.000  -0.4429   0.02686   0.02012  -0.0496   0.9554   0.0174
  -5.750  -0.4167   0.02351   0.01628  -0.0500   0.9524   0.0181
  -5.500  -0.3930   0.02151   0.01404  -0.0497   0.9474   0.0190
  -5.000  -0.3368   0.01952   0.01171  -0.0505   0.9403   0.0228
  -4.750  -0.3078   0.01820   0.01020  -0.0508   0.9373   0.0238
  -4.500  -0.2829   0.01717   0.00905  -0.0502   0.9322   0.0250
  -4.250  -0.2559   0.01630   0.00807  -0.0501   0.9285   0.0262
  -4.000  -0.2277   0.01574   0.00744  -0.0503   0.9255   0.0276
  -3.750  -0.2060   0.01466   0.00632  -0.0495   0.9209   0.0311
  -3.500  -0.1818   0.01408   0.00570  -0.0491   0.9164   0.0335
  -3.250  -0.1556   0.01359   0.00513  -0.0489   0.9128   0.0366
  -3.000  -0.1294   0.01323   0.00466  -0.0488   0.9093   0.0401
  -2.750  -0.1056   0.01284   0.00426  -0.0482   0.9046   0.0497
  -2.500  -0.0800   0.01246   0.00392  -0.0480   0.9010   0.0703
  -2.250  -0.0654   0.01051   0.00348  -0.0469   0.8973   0.4424
  -2.000  -0.0552   0.00978   0.00395  -0.0426   0.8921   0.7480
  -1.750  -0.0342   0.00997   0.00422  -0.0404   0.8882   0.8035
  -1.500  -0.0099   0.01010   0.00430  -0.0391   0.8852   0.8279
  -1.250   0.0150   0.01022   0.00440  -0.0380   0.8829   0.8445
  -1.000   0.0353   0.01042   0.00460  -0.0360   0.8784   0.8618
  -0.750   0.0591   0.01051   0.00465  -0.0349   0.8747   0.8727
  -0.500   0.0866   0.01050   0.00459  -0.0349   0.8718   0.8766
  -0.250   0.1149   0.01048   0.00452  -0.0351   0.8696   0.8807
   0.000   0.1407   0.01051   0.00454  -0.0350   0.8659   0.8843
   0.250   0.1668   0.01053   0.00457  -0.0347   0.8621   0.8877
   0.500   0.1940   0.01054   0.00458  -0.0347   0.8590   0.8916
   0.750   0.2221   0.01053   0.00457  -0.0349   0.8563   0.8956
   1.000   0.2480   0.01054   0.00462  -0.0346   0.8514   0.8990
   1.250   0.2742   0.01054   0.00466  -0.0343   0.8463   0.9027
   1.500   0.3022   0.01051   0.00467  -0.0343   0.8424   0.9069
   1.750   0.3277   0.01052   0.00475  -0.0340   0.8367   0.9116
   2.000   0.3553   0.01040   0.00467  -0.0336   0.8286   0.9147
   2.250   0.3799   0.01010   0.00441  -0.0323   0.8079   0.9189
   2.500   0.4046   0.00983   0.00413  -0.0310   0.7824   0.9238
   2.750   0.4290   0.00965   0.00394  -0.0297   0.7477   0.9282
   3.000   0.4508   0.00957   0.00360  -0.0277   0.6503   0.9332
   3.250   0.4539   0.01092   0.00364  -0.0227   0.3847   0.9415
   3.500   0.4605   0.01274   0.00427  -0.0198   0.1209   0.9506
   3.750   0.4827   0.01361   0.00478  -0.0194   0.0512   0.9570
   4.000   0.5092   0.01407   0.00526  -0.0195   0.0399   0.9637
   4.250   0.5374   0.01470   0.00594  -0.0200   0.0340   0.9695
   4.750   0.5949   0.01601   0.00749  -0.0213   0.0286   0.9827
   5.000   0.6234   0.01686   0.00840  -0.0219   0.0267   0.9914
   5.250   0.6469   0.01781   0.00937  -0.0217   0.0245   1.0000
   5.500   0.6652   0.01915   0.01073  -0.0203   0.0227   1.0000
   5.750   0.6881   0.02007   0.01175  -0.0197   0.0219   1.0000
   6.000   0.7126   0.02125   0.01309  -0.0192   0.0211   1.0000
   6.250   0.7378   0.02268   0.01468  -0.0188   0.0204   1.0000
   6.500   0.7631   0.02439   0.01662  -0.0184   0.0197   1.0000
   6.750   0.7869   0.02591   0.01837  -0.0178   0.0184   1.0000
   7.000   0.8091   0.02702   0.01964  -0.0173   0.0169   1.0000
   7.250   0.8303   0.02876   0.02163  -0.0165   0.0161   1.0000
   7.500   0.8495   0.03077   0.02388  -0.0156   0.0155   1.0000
   7.750   0.8657   0.03349   0.02694  -0.0143   0.0150   1.0000
   8.000   0.8743   0.03786   0.03176  -0.0122   0.0145   1.0000
   8.250   0.8845   0.04066   0.03502  -0.0099   0.0139   1.0000
   8.500   0.8931   0.04411   0.03897  -0.0074   0.0126   1.0000
   8.750   0.8943   0.04839   0.04363  -0.0047   0.0124   1.0000
   9.000   0.8912   0.05276   0.04834  -0.0020   0.0122   1.0000
   9.250   0.8840   0.05709   0.05295   0.0006   0.0121   1.0000
   9.500   0.8720   0.06108   0.05717   0.0033   0.0120   1.0000
   9.750   0.8552   0.06493   0.06118   0.0059   0.0120   1.0000
  10.000   0.8369   0.06909   0.06549   0.0071   0.0121   1.0000
  10.250   0.8177   0.07390   0.07042   0.0067   0.0121   1.0000
  10.500   0.7990   0.07944   0.07606   0.0046   0.0123   1.0000
  10.750   0.7805   0.08664   0.08334   0.0001   0.0125   1.0000
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