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CHEN AIRFOIL (chen-il) Xfoil prediction polar at RE=500,000 Ncrit=5


Details Polar file
Airfoil: CHEN AIRFOIL (chen-il)
Reynolds number: 500,000
Max Cl/Cd: 56.54 at α=12°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-chen-il-500000-n5.txt
Download as CSV file: xf-chen-il-500000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: CHEN AIRFOIL                                    
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.500 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.000   0.0118   0.09620   0.09083  -0.0617   0.3239   0.0132
  -8.750   0.0190   0.09364   0.08828  -0.0633   0.3238   0.0154
  -8.500   0.0250   0.09104   0.08570  -0.0646   0.3237   0.0155
  -8.000   0.0357   0.08598   0.08068  -0.0669   0.3237   0.0155
  -7.750   0.0388   0.08336   0.07809  -0.0681   0.3237   0.0156
  -7.250   0.0391   0.07839   0.07319  -0.0699   0.3235   0.0156
  -7.000   0.0394   0.07632   0.07113  -0.0692   0.3234   0.0156
  -6.750   0.0364   0.07444   0.06927  -0.0678   0.3233   0.0156
  -6.250   0.0439   0.06999   0.06484  -0.0668   0.3233   0.0156
  -5.750   0.0605   0.06650   0.06136  -0.0651   0.3232   0.0148
  -5.500   0.0660   0.06423   0.05908  -0.0644   0.3231   0.0142
  -5.250   0.0705   0.06144   0.05626  -0.0636   0.3230   0.0133
  -5.000   0.0794   0.05928   0.05407  -0.0628   0.3229   0.0136
  -4.750   0.0884   0.05689   0.05163  -0.0619   0.3228   0.0135
  -4.500   0.0983   0.05447   0.04916  -0.0607   0.3229   0.0135
  -4.250   0.1103   0.05234   0.04697  -0.0595   0.3227   0.0138
  -4.000   0.1233   0.05029   0.04486  -0.0582   0.3227   0.0141
  -3.750   0.1363   0.04804   0.04253  -0.0567   0.3226   0.0143
  -3.500   0.1499   0.04579   0.04018  -0.0549   0.3225   0.0146
  -3.250   0.1608   0.04241   0.03663  -0.0522   0.3224   0.0154
  -3.000   0.1722   0.03941   0.03348  -0.0493   0.3224   0.0156
  -2.750   0.1889   0.03797   0.03196  -0.0475   0.3223   0.0159
  -2.500   0.2055   0.03633   0.03022  -0.0454   0.3222   0.0161
  -2.000   0.2078   0.01944   0.01150  -0.0324   0.3221   0.0191
  -1.750   0.2355   0.01874   0.01065  -0.0322   0.3220   0.0194
  -1.500   0.2625   0.01838   0.01022  -0.0318   0.3219   0.0200
  -1.250   0.2898   0.01812   0.00992  -0.0315   0.3217   0.0207
  -1.000   0.3169   0.01792   0.00967  -0.0312   0.3217   0.0215
  -0.750   0.3440   0.01776   0.00945  -0.0308   0.3216   0.0224
  -0.250   0.3970   0.01750   0.00914  -0.0299   0.3213   0.0244
   0.000   0.4227   0.01749   0.00912  -0.0293   0.3212   0.0264
   0.250   0.4480   0.01746   0.00908  -0.0287   0.3211   0.0284
   0.500   0.4729   0.01750   0.00914  -0.0279   0.3210   0.0310
   0.750   0.4973   0.01754   0.00917  -0.0271   0.3209   0.0338
   1.000   0.5216   0.01761   0.00926  -0.0262   0.3207   0.0374
   1.250   0.5455   0.01771   0.00934  -0.0253   0.3207   0.0409
   1.500   0.5695   0.01787   0.00953  -0.0244   0.3206   0.0452
   1.750   0.5932   0.01798   0.00964  -0.0235   0.3204   0.0485
   2.000   0.6170   0.01814   0.00983  -0.0226   0.3203   0.0518
   2.250   0.6405   0.01833   0.01002  -0.0217   0.3202   0.0553
   2.500   0.6639   0.01852   0.01025  -0.0207   0.3201   0.0591
   2.750   0.6876   0.01875   0.01049  -0.0199   0.3200   0.0630
   3.000   0.7110   0.01893   0.01066  -0.0189   0.3198   0.0662
   3.250   0.7341   0.01907   0.01082  -0.0180   0.3195   0.0695
   3.500   0.7575   0.01926   0.01102  -0.0171   0.3193   0.0730
   3.750   0.7806   0.01949   0.01126  -0.0162   0.3191   0.0766
   4.000   0.8033   0.01975   0.01153  -0.0152   0.3190   0.0798
   4.250   0.8259   0.02000   0.01182  -0.0143   0.3188   0.0844
   4.500   0.8484   0.02031   0.01215  -0.0133   0.3187   0.0881
   4.750   0.8710   0.02063   0.01248  -0.0124   0.3185   0.0911
   5.000   0.8929   0.02098   0.01287  -0.0114   0.3183   0.0955
   5.250   0.9147   0.02139   0.01334  -0.0105   0.3181   0.1009
   5.750   0.9561   0.02236   0.01442  -0.0083   0.3173   0.1138
   6.000   0.9770   0.02263   0.01477  -0.0072   0.3172   0.1224
   6.250   0.9973   0.02299   0.01520  -0.0060   0.3170   0.1352
   6.500   1.0171   0.02332   0.01561  -0.0047   0.3167   0.1475
   6.750   1.0365   0.02368   0.01605  -0.0034   0.3165   0.1571
   7.250   1.0727   0.02449   0.01703  -0.0006   0.3155   0.1771
   7.500   1.0900   0.02503   0.01766   0.0009   0.3151   0.1869
   7.750   1.1060   0.02559   0.01834   0.0025   0.3146   0.2097
   8.250   1.2714   0.02690   0.02137  -0.0239   0.3131   0.9846
   8.750   1.3545   0.02857   0.02320  -0.0319   0.3115   0.9945
   9.000   1.3894   0.02925   0.02394  -0.0345   0.3108   0.9975
   9.250   1.4161   0.02997   0.02474  -0.0355   0.3103   0.9992
   9.500   1.4380   0.03071   0.02555  -0.0355   0.3097   1.0000
   9.750   1.4515   0.03127   0.02616  -0.0336   0.3092   1.0000
  10.000   1.4766   0.03091   0.02579  -0.0330   0.3086   1.0000
  10.250   1.4824   0.03202   0.02698  -0.0303   0.3082   1.0000
  10.500   1.5067   0.03170   0.02666  -0.0296   0.3077   1.0000
  10.750   1.5406   0.03068   0.02557  -0.0300   0.3071   1.0000
  11.000   1.5843   0.02900   0.02377  -0.0317   0.3063   1.0000
  11.250   1.5860   0.03041   0.02531  -0.0285   0.3063   1.0000
  11.500   1.6224   0.02935   0.02417  -0.0293   0.3057   1.0000
  11.750   1.6441   0.02933   0.02416  -0.0285   0.3053   1.0000
  12.000   1.6584   0.02933   0.02422  -0.0265   0.3041   1.0000
  12.500   1.0901   0.08704   0.08289  -0.0191   0.2606   1.0000
  13.000   1.1197   0.08847   0.08434  -0.0184   0.2593   1.0000
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