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NACA 64A410 (naca64a410-il) Xfoil prediction polar at RE=500,000 Ncrit=9


Details Polar file
Airfoil: NACA 64A410 (naca64a410-il)
Reynolds number: 500,000
Max Cl/Cd: 100.72 at α=3.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-naca64a410-il-500000.txt
Download as CSV file: xf-naca64a410-il-500000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NACA 64A410                                     
  
 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
  ------ -------- --------- --------- -------- -------- --------
  -9.250  -0.4386   0.08360   0.08140  -0.0473   1.0000   0.0186
  -9.000  -0.4481   0.07985   0.07770  -0.0475   1.0000   0.0189
  -8.750  -0.4686   0.07771   0.07563  -0.0449   1.0000   0.0189
  -8.500  -0.4779   0.07386   0.07182  -0.0471   0.9978   0.0190
  -8.250  -0.4609   0.06567   0.06356  -0.0610   0.9928   0.0193
  -8.000  -0.4448   0.06049   0.05830  -0.0679   0.9883   0.0197
  -7.750  -0.4242   0.05557   0.05326  -0.0746   0.9852   0.0205
  -7.500  -0.4023   0.05033   0.04784  -0.0807   0.9821   0.0218
  -7.250  -0.3744   0.04526   0.04230  -0.0851   0.9766   0.0248
  -6.250  -0.2957   0.02083   0.01578  -0.0919   0.9595   0.0194
  -6.000  -0.2669   0.01886   0.01355  -0.0926   0.9568   0.0198
  -5.750  -0.2419   0.01839   0.01305  -0.0925   0.9521   0.0213
  -5.500  -0.2164   0.01686   0.01129  -0.0920   0.9473   0.0219
  -5.250  -0.1889   0.01542   0.00964  -0.0919   0.9437   0.0225
  -5.000  -0.1627   0.01439   0.00847  -0.0915   0.9397   0.0233
  -4.750  -0.1380   0.01361   0.00758  -0.0908   0.9341   0.0241
  -4.500  -0.1120   0.01277   0.00664  -0.0904   0.9298   0.0256
  -4.250  -0.0876   0.01179   0.00562  -0.0898   0.9252   0.0273
  -4.000  -0.0627   0.01129   0.00510  -0.0892   0.9193   0.0290
  -3.750  -0.0366   0.01083   0.00458  -0.0887   0.9140   0.0312
  -3.500  -0.0111   0.01055   0.00425  -0.0881   0.9071   0.0335
  -3.250   0.0138   0.00986   0.00351  -0.0875   0.9005   0.0387
  -3.000   0.0400   0.00957   0.00317  -0.0871   0.8942   0.0443
  -2.750   0.0662   0.00912   0.00266  -0.0867   0.8880   0.0565
  -2.500   0.0928   0.00873   0.00246  -0.0865   0.8826   0.1148
  -2.250   0.1191   0.00847   0.00228  -0.0863   0.8755   0.1447
  -2.000   0.1439   0.00769   0.00205  -0.0862   0.8698   0.3170
  -1.750   0.1673   0.00698   0.00199  -0.0857   0.8619   0.5197
  -1.500   0.1937   0.00679   0.00193  -0.0853   0.8552   0.5833
  -1.250   0.2197   0.00666   0.00192  -0.0849   0.8463   0.6277
  -1.000   0.2465   0.00659   0.00185  -0.0845   0.8375   0.6544
  -0.750   0.2730   0.00651   0.00182  -0.0841   0.8280   0.6845
  -0.500   0.2995   0.00645   0.00182  -0.0837   0.8188   0.7120
  -0.250   0.3265   0.00642   0.00179  -0.0834   0.8104   0.7323
   0.000   0.3532   0.00638   0.00178  -0.0831   0.8003   0.7532
   0.250   0.3801   0.00636   0.00176  -0.0828   0.7900   0.7690
   0.500   0.4071   0.00634   0.00173  -0.0826   0.7800   0.7830
   1.000   0.4604   0.00629   0.00172  -0.0819   0.7585   0.8140
   1.250   0.4864   0.00625   0.00174  -0.0815   0.7476   0.8345
   1.500   0.5115   0.00620   0.00177  -0.0807   0.7357   0.8654
   1.750   0.5413   0.00612   0.00182  -0.0809   0.7221   0.9332
   2.000   0.5833   0.00619   0.00182  -0.0840   0.7009   1.0000
   2.250   0.6089   0.00632   0.00186  -0.0836   0.6810   1.0000
   2.500   0.6345   0.00645   0.00193  -0.0832   0.6587   1.0000
   2.750   0.6597   0.00662   0.00201  -0.0827   0.6344   1.0000
   3.000   0.6846   0.00681   0.00210  -0.0822   0.6097   1.0000
   3.250   0.7091   0.00704   0.00222  -0.0816   0.5790   1.0000
   3.500   0.7327   0.00735   0.00239  -0.0808   0.5399   1.0000
   3.750   0.7544   0.00780   0.00259  -0.0797   0.4799   1.0000
   4.000   0.7673   0.00906   0.00308  -0.0774   0.3226   1.0000
   4.250   0.7809   0.01047   0.00374  -0.0754   0.1690   1.0000
   4.500   0.8025   0.01110   0.00415  -0.0746   0.1287   1.0000
   4.750   0.8265   0.01150   0.00449  -0.0741   0.1146   1.0000
   5.000   0.8509   0.01185   0.00481  -0.0736   0.0972   1.0000
   5.250   0.8717   0.01257   0.00527  -0.0726   0.0491   1.0000
   5.500   0.8931   0.01323   0.00595  -0.0716   0.0372   1.0000
   5.750   0.9144   0.01391   0.00664  -0.0705   0.0300   1.0000
   6.000   0.9361   0.01452   0.00730  -0.0695   0.0270   1.0000
   6.250   0.9580   0.01508   0.00790  -0.0686   0.0248   1.0000
   6.500   0.9788   0.01574   0.00859  -0.0675   0.0231   1.0000
   6.750   0.9935   0.01701   0.00994  -0.0654   0.0213   1.0000
   7.000   1.0135   0.01776   0.01075  -0.0642   0.0205   1.0000
   7.250   1.0334   0.01854   0.01160  -0.0630   0.0197   1.0000
   7.500   1.0526   0.01946   0.01259  -0.0617   0.0191   1.0000
   7.750   1.0719   0.02046   0.01365  -0.0604   0.0184   1.0000
   8.000   1.0917   0.02152   0.01480  -0.0592   0.0179   1.0000
   8.250   1.1119   0.02266   0.01601  -0.0581   0.0173   1.0000
   8.500   1.1322   0.02380   0.01722  -0.0571   0.0168   1.0000
   8.750   1.1521   0.02515   0.01862  -0.0562   0.0160   1.0000
   9.000   1.1748   0.02910   0.02278  -0.0559   0.0153   1.0000
   9.250   1.1928   0.03102   0.02493  -0.0545   0.0151   1.0000
   9.500   1.2090   0.03287   0.02704  -0.0529   0.0149   1.0000
   9.750   1.2224   0.03527   0.02973  -0.0510   0.0148   1.0000
  10.000   1.2316   0.03824   0.03301  -0.0487   0.0147   1.0000
  10.250   1.2367   0.04121   0.03629  -0.0459   0.0146   1.0000
  10.500   1.2375   0.04410   0.03950  -0.0426   0.0143   1.0000
  10.750   1.2319   0.04714   0.04284  -0.0388   0.0141   1.0000
  11.000   1.2192   0.05053   0.04650  -0.0345   0.0142   1.0000
  11.250   1.2013   0.05428   0.05049  -0.0304   0.0144   1.0000
  11.500   1.1803   0.05845   0.05489  -0.0272   0.0146   1.0000
  11.750   1.0804   0.05237   0.04904  -0.0141   0.0163   1.0000
  12.000   1.0656   0.05592   0.05269  -0.0124   0.0161   1.0000
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