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NACA 66-210 (naca66210-il) Xfoil prediction polar at RE=500,000 Ncrit=0


Details Polar file
Airfoil: NACA 66-210 (naca66210-il)
Reynolds number: 500,000
Max Cl/Cd: 67.52 at α=3°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-naca66210-il-500000.txt
Download as CSV file: xf-naca66210-il-500000.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 =     0.500 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -13.000  -0.5264   0.13599   0.13356  -0.0210   1.0000   0.0135
 -12.750  -0.5208   0.13275   0.13033  -0.0215   1.0000   0.0138
 -12.500  -0.5152   0.12951   0.12707  -0.0223   1.0000   0.0141
 -12.250  -0.5101   0.12634   0.12391  -0.0232   1.0000   0.0147
 -12.000  -0.5071   0.12255   0.12012  -0.0247   1.0000   0.0149
 -11.750  -0.5042   0.11880   0.11638  -0.0262   1.0000   0.0153
 -11.500  -0.5023   0.11481   0.11241  -0.0279   1.0000   0.0157
  -8.500  -0.5593   0.06045   0.05788  -0.0545   0.9904   0.0192
  -8.250  -0.5488   0.05511   0.05226  -0.0578   0.9859   0.0192
  -8.000  -0.5580   0.04432   0.04105  -0.0605   0.9782   0.0200
  -7.750  -0.5410   0.04043   0.03702  -0.0623   0.9740   0.0208
  -7.500  -0.5221   0.03784   0.03429  -0.0631   0.9682   0.0215
  -7.250  -0.5011   0.03542   0.03168  -0.0639   0.9635   0.0227
  -7.000  -0.4843   0.03313   0.02916  -0.0630   0.9558   0.0245
  -6.750  -0.4525   0.03457   0.03034  -0.0623   0.9514   0.0283
  -6.000  -0.4162   0.02053   0.01498  -0.0558   0.9299   0.0231
  -5.750  -0.3930   0.01851   0.01268  -0.0547   0.9254   0.0232
  -5.500  -0.3686   0.01734   0.01133  -0.0541   0.9214   0.0243
  -5.250  -0.3438   0.01646   0.01034  -0.0535   0.9167   0.0258
  -5.000  -0.3184   0.01538   0.00912  -0.0530   0.9128   0.0264
  -4.750  -0.2931   0.01451   0.00813  -0.0524   0.9095   0.0271
  -4.500  -0.2680   0.01384   0.00740  -0.0520   0.9052   0.0278
  -4.250  -0.2430   0.01331   0.00681  -0.0515   0.9010   0.0289
  -4.000  -0.2221   0.01197   0.00540  -0.0503   0.8972   0.0307
  -3.750  -0.1991   0.01133   0.00470  -0.0496   0.8938   0.0326
  -3.500  -0.1748   0.01090   0.00424  -0.0491   0.8897   0.0351
  -3.250  -0.1497   0.01055   0.00384  -0.0487   0.8862   0.0381
  -3.000  -0.1237   0.01030   0.00355  -0.0484   0.8832   0.0409
  -2.750  -0.0986   0.00992   0.00313  -0.0480   0.8805   0.0496
  -2.500  -0.0734   0.00958   0.00288  -0.0477   0.8767   0.0752
  -2.250  -0.0633   0.00732   0.00238  -0.0459   0.8722   0.5443
  -2.000  -0.0432   0.00684   0.00253  -0.0443   0.8692   0.7324
  -1.750  -0.0169   0.00691   0.00262  -0.0437   0.8670   0.7680
  -1.500   0.0091   0.00701   0.00276  -0.0432   0.8641   0.7897
  -1.250   0.0338   0.00719   0.00300  -0.0422   0.8610   0.8141
  -1.000   0.0597   0.00731   0.00313  -0.0415   0.8581   0.8274
  -0.750   0.0868   0.00738   0.00319  -0.0413   0.8556   0.8348
  -0.500   0.1140   0.00744   0.00323  -0.0411   0.8535   0.8416
  -0.250   0.1399   0.00759   0.00339  -0.0406   0.8512   0.8519
   0.000   0.1660   0.00764   0.00347  -0.0402   0.8478   0.8569
   0.250   0.1934   0.00763   0.00347  -0.0402   0.8440   0.8599
   0.500   0.2217   0.00760   0.00341  -0.0403   0.8404   0.8626
   0.750   0.2497   0.00757   0.00338  -0.0404   0.8360   0.8656
   1.000   0.2765   0.00746   0.00325  -0.0402   0.8278   0.8684
   1.250   0.3033   0.00738   0.00318  -0.0398   0.8203   0.8708
   1.500   0.3304   0.00731   0.00313  -0.0396   0.8130   0.8734
   1.750   0.3573   0.00724   0.00307  -0.0393   0.8043   0.8764
   2.000   0.3839   0.00712   0.00294  -0.0389   0.7907   0.8798
   2.250   0.4095   0.00700   0.00281  -0.0382   0.7728   0.8829
   2.500   0.4340   0.00692   0.00270  -0.0372   0.7476   0.8858
   2.750   0.4585   0.00691   0.00265  -0.0364   0.7158   0.8892
   3.000   0.4794   0.00710   0.00260  -0.0348   0.6299   0.8933
   3.250   0.4786   0.00886   0.00307  -0.0297   0.3456   0.8987
   3.500   0.4820   0.01074   0.00374  -0.0260   0.0715   0.9046
   3.750   0.5045   0.01120   0.00409  -0.0252   0.0463   0.9096
   4.000   0.5258   0.01154   0.00446  -0.0240   0.0388   0.9139
   4.250   0.5469   0.01199   0.00499  -0.0227   0.0348   0.9190
   4.750   0.5890   0.01279   0.00591  -0.0203   0.0295   0.9297
   5.000   0.6081   0.01337   0.00653  -0.0187   0.0274   0.9364
   5.250   0.6237   0.01424   0.00747  -0.0165   0.0259   0.9435
   5.500   0.6401   0.01573   0.00905  -0.0145   0.0248   0.9516
   5.750   0.6621   0.01640   0.00981  -0.0134   0.0241   0.9592
   6.000   0.6872   0.01712   0.01060  -0.0130   0.0237   0.9670
   6.250   0.7163   0.01819   0.01177  -0.0134   0.0232   0.9736
   6.500   0.7483   0.01952   0.01323  -0.0145   0.0226   0.9790
   6.750   0.7809   0.02117   0.01506  -0.0156   0.0221   0.9846
   7.000   0.8136   0.02272   0.01680  -0.0169   0.0210   0.9912
   7.250   0.8419   0.02471   0.01903  -0.0174   0.0204   1.0000
   7.500   0.8621   0.02786   0.02254  -0.0162   0.0207   1.0000
   8.500   0.8356   0.03847   0.03497  -0.0025   0.0244   1.0000
   8.750   0.8349   0.04239   0.03913  -0.0001   0.0244   1.0000
   9.000   0.8536   0.04111   0.03811   0.0021   0.0220   1.0000
   9.250   0.8487   0.04470   0.04193   0.0048   0.0208   1.0000
   9.500   0.8340   0.04820   0.04558   0.0083   0.0205   1.0000
   9.750   0.8153   0.05158   0.04910   0.0113   0.0202   1.0000
  10.000   0.7929   0.05557   0.05322   0.0133   0.0201   1.0000
  10.250   0.7667   0.06046   0.05822   0.0141   0.0203   1.0000
  10.500   0.7368   0.06658   0.06447   0.0133   0.0209   1.0000
  10.750   0.7058   0.07399   0.07197   0.0104   0.0217   1.0000
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