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NACA 65-206 (naca65206-il) Xfoil prediction polar at RE=500,000 Ncrit=9


Details Polar file
Airfoil: NACA 65-206 (naca65206-il)
Reynolds number: 500,000
Max Cl/Cd: 71.17 at α=2°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-naca65206-il-500000.txt
Download as CSV file: xf-naca65206-il-500000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NACA 65-206                                     
  
 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
  ------ -------- --------- --------- -------- -------- --------
  -8.250  -0.5276   0.08792   0.08565  -0.0140   1.0000   0.0103
  -8.000  -0.5266   0.08407   0.08182  -0.0164   1.0000   0.0103
  -7.750  -0.5282   0.08008   0.07787  -0.0196   1.0000   0.0103
  -7.500  -0.5243   0.07531   0.07308  -0.0247   1.0000   0.0103
  -7.250  -0.5183   0.07075   0.06848  -0.0280   1.0000   0.0103
  -7.000  -0.5107   0.06633   0.06399  -0.0305   1.0000   0.0103
  -6.750  -0.5016   0.06195   0.05953  -0.0325   1.0000   0.0103
  -6.500  -0.4911   0.05758   0.05505  -0.0339   1.0000   0.0104
  -6.250  -0.4793   0.05337   0.05071  -0.0347   1.0000   0.0104
  -6.000  -0.4665   0.04912   0.04629  -0.0351   1.0000   0.0104
  -5.750  -0.4426   0.02654   0.02367  -0.0356   1.0000   0.0113
  -5.500  -0.4324   0.02353   0.02057  -0.0352   1.0000   0.0122
  -5.250  -0.4194   0.02116   0.01808  -0.0345   1.0000   0.0133
  -5.000  -0.4058   0.01889   0.01564  -0.0333   1.0000   0.0149
  -4.750  -0.3821   0.02002   0.01659  -0.0307   1.0000   0.0194
  -4.500  -0.3649   0.01792   0.01424  -0.0297   0.9996   0.0196
  -4.250  -0.3331   0.01484   0.01082  -0.0316   0.9971   0.0197
  -4.000  -0.3005   0.01192   0.00755  -0.0335   0.9948   0.0197
  -3.500  -0.2672   0.01684   0.01150  -0.0327   0.9946   0.0175
  -3.250  -0.2336   0.01359   0.00784  -0.0331   0.9926   0.0149
  -3.000  -0.1999   0.01212   0.00620  -0.0340   0.9896   0.0169
  -2.750  -0.1634   0.01227   0.00628  -0.0359   0.9863   0.0198
  -2.500  -0.1306   0.01002   0.00393  -0.0370   0.9841   0.0213
  -2.250  -0.0942   0.00913   0.00299  -0.0391   0.9820   0.0260
  -2.000  -0.0606   0.00874   0.00253  -0.0404   0.9775   0.0304
  -1.750  -0.0250   0.00818   0.00190  -0.0422   0.9738   0.0394
  -1.500   0.0058   0.00585   0.00150  -0.0444   0.9704   0.6111
  -1.250   0.0342   0.00536   0.00157  -0.0444   0.9649   0.7765
  -1.000   0.0606   0.00519   0.00164  -0.0435   0.9580   0.8546
  -0.750   0.0853   0.00514   0.00166  -0.0423   0.9501   0.8935
  -0.500   0.1095   0.00510   0.00164  -0.0409   0.9422   0.9195
  -0.250   0.1325   0.00505   0.00160  -0.0395   0.9321   0.9371
   0.000   0.1581   0.00501   0.00154  -0.0388   0.9229   0.9510
   0.500   0.2198   0.00493   0.00145  -0.0398   0.9050   0.9751
   0.750   0.2555   0.00491   0.00142  -0.0416   0.8959   0.9848
   1.250   0.3176   0.00489   0.00130  -0.0427   0.8531   1.0000
   1.500   0.3396   0.00495   0.00126  -0.0413   0.8220   1.0000
   1.750   0.3610   0.00513   0.00119  -0.0397   0.7578   1.0000
   2.000   0.3843   0.00540   0.00125  -0.0386   0.6906   1.0000
   2.250   0.4042   0.00614   0.00135  -0.0371   0.5276   1.0000
   2.500   0.4159   0.00888   0.00210  -0.0355   0.0435   1.0000
   2.750   0.4421   0.00936   0.00260  -0.0352   0.0307   1.0000
   3.000   0.4671   0.01014   0.00345  -0.0347   0.0243   1.0000
   3.250   0.4914   0.01104   0.00443  -0.0340   0.0225   1.0000
   3.500   0.5168   0.01175   0.00524  -0.0335   0.0214   1.0000
   3.750   0.5424   0.01242   0.00597  -0.0331   0.0191   1.0000
   4.000   0.5675   0.01346   0.00708  -0.0325   0.0173   1.0000
   4.250   0.5929   0.01495   0.00868  -0.0318   0.0161   1.0000
   4.500   0.6191   0.01688   0.01077  -0.0310   0.0159   1.0000
   5.000   0.6678   0.02149   0.01593  -0.0296   0.0119   1.0000
   5.250   0.6911   0.02501   0.01988  -0.0281   0.0117   1.0000
   5.500   0.7148   0.02794   0.02309  -0.0270   0.0141   1.0000
   5.750   0.7339   0.03169   0.02718  -0.0258   0.0138   1.0000
   6.000   0.7384   0.04070   0.03677  -0.0239   0.0129   1.0000
   6.250   0.7535   0.04463   0.04098  -0.0228   0.0129   1.0000
   6.500   0.7668   0.04868   0.04531  -0.0218   0.0129   1.0000
   6.750   0.7781   0.05288   0.04975  -0.0210   0.0129   1.0000
   7.000   0.7877   0.05707   0.05416  -0.0203   0.0128   1.0000
   7.250   0.7950   0.06138   0.05867  -0.0197   0.0128   1.0000
   7.500   0.7999   0.06573   0.06320  -0.0195   0.0128   1.0000
   7.750   0.8022   0.07009   0.06771  -0.0195   0.0128   1.0000
   8.000   0.8016   0.07448   0.07221  -0.0198   0.0127   1.0000
   8.250   0.7974   0.07886   0.07669  -0.0205   0.0127   1.0000
   8.500   0.7871   0.08275   0.08065  -0.0208   0.0127   1.0000
   8.750   0.7766   0.08753   0.08547  -0.0236   0.0127   1.0000
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