Airfoil Tools
Search 1636 airfoils Google+
You have 0 airfoils loaded.
Your Reynold number range is 50,000 to 1,000,000. (set)

NACA 63-415 AIRFOIL (n63415-il) Xfoil prediction polar at RE=500,000 Ncrit=9


Details Polar file
Airfoil: NACA 63-415 AIRFOIL (n63415-il)
Reynolds number: 500,000
Max Cl/Cd: 100.89 at α=6°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-n63415-il-500000.txt
Download as CSV file: xf-n63415-il-500000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NACA 63-415 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 =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -15.500  -0.7187   0.09149   0.08869  -0.0506   1.0000   0.0203
 -15.250  -0.7729   0.07691   0.07384  -0.0605   1.0000   0.0198
 -15.000  -0.8084   0.06743   0.06413  -0.0667   1.0000   0.0197
 -14.750  -0.8319   0.06085   0.05736  -0.0706   1.0000   0.0197
 -14.500  -0.8540   0.05497   0.05126  -0.0737   1.0000   0.0197
 -14.250  -0.8685   0.05044   0.04655  -0.0757   1.0000   0.0198
 -14.000  -0.8800   0.04651   0.04243  -0.0770   1.0000   0.0199
 -13.750  -0.8880   0.04322   0.03896  -0.0778   1.0000   0.0202
 -13.500  -0.8935   0.04030   0.03587  -0.0781   1.0000   0.0204
 -13.250  -0.8968   0.03782   0.03321  -0.0780   1.0000   0.0206
 -13.000  -0.8983   0.03561   0.03083  -0.0776   1.0000   0.0209
 -12.750  -0.8990   0.03389   0.02892  -0.0768   1.0000   0.0211
 -12.500  -0.8982   0.03149   0.02639  -0.0755   1.0000   0.0215
 -12.250  -0.8916   0.03012   0.02500  -0.0742   1.0000   0.0219
 -12.000  -0.8861   0.02905   0.02391  -0.0726   1.0000   0.0223
 -11.750  -0.8841   0.02816   0.02299  -0.0702   1.0000   0.0226
 -11.500  -0.8881   0.02742   0.02222  -0.0666   1.0000   0.0229
 -11.250  -0.8784   0.02661   0.02135  -0.0652   0.9980   0.0235
 -11.000  -0.8472   0.02529   0.01989  -0.0676   0.9906   0.0243
 -10.750  -0.8150   0.02412   0.01855  -0.0700   0.9841   0.0251
 -10.500  -0.7848   0.02250   0.01682  -0.0720   0.9774   0.0259
 -10.250  -0.7513   0.02138   0.01571  -0.0744   0.9710   0.0269
 -10.000  -0.7183   0.02055   0.01482  -0.0766   0.9619   0.0279
  -9.750  -0.6847   0.01981   0.01400  -0.0786   0.9525   0.0293
  -9.500  -0.6548   0.01923   0.01327  -0.0797   0.9401   0.0303
  -9.250  -0.6348   0.01799   0.01199  -0.0793   0.9258   0.0316
  -9.000  -0.6136   0.01743   0.01139  -0.0787   0.9125   0.0328
  -8.750  -0.5921   0.01694   0.01082  -0.0780   0.9009   0.0341
  -8.500  -0.5700   0.01651   0.01028  -0.0772   0.8906   0.0355
  -8.250  -0.5504   0.01579   0.00949  -0.0763   0.8798   0.0370
  -8.000  -0.5291   0.01523   0.00890  -0.0757   0.8707   0.0388
  -7.750  -0.5059   0.01481   0.00843  -0.0752   0.8614   0.0406
  -7.500  -0.4818   0.01445   0.00798  -0.0748   0.8534   0.0423
  -7.250  -0.4598   0.01379   0.00727  -0.0742   0.8447   0.0446
  -7.000  -0.4354   0.01339   0.00683  -0.0739   0.8372   0.0472
  -6.750  -0.4100   0.01307   0.00645  -0.0737   0.8292   0.0498
  -6.500  -0.3854   0.01260   0.00590  -0.0734   0.8223   0.0530
  -6.250  -0.3596   0.01222   0.00553  -0.0734   0.8144   0.0572
  -6.000  -0.3336   0.01187   0.00511  -0.0732   0.8077   0.0623
  -5.750  -0.3072   0.01151   0.00475  -0.0732   0.8005   0.0702
  -5.500  -0.2813   0.01100   0.00433  -0.0732   0.7934   0.0903
  -5.250  -0.2557   0.01038   0.00392  -0.0734   0.7870   0.1418
  -5.000  -0.2301   0.00962   0.00351  -0.0737   0.7799   0.2211
  -4.750  -0.2041   0.00888   0.00313  -0.0741   0.7736   0.3201
  -4.500  -0.1767   0.00844   0.00296  -0.0745   0.7668   0.3958
  -4.250  -0.1483   0.00825   0.00284  -0.0747   0.7604   0.4390
  -4.000  -0.1198   0.00816   0.00277  -0.0749   0.7543   0.4677
  -3.750  -0.0909   0.00809   0.00271  -0.0751   0.7474   0.4880
  -3.500  -0.0620   0.00806   0.00264  -0.0753   0.7414   0.5044
  -3.250  -0.0331   0.00804   0.00260  -0.0755   0.7352   0.5192
  -3.000  -0.0042   0.00802   0.00258  -0.0756   0.7286   0.5353
  -2.750   0.0246   0.00805   0.00257  -0.0758   0.7227   0.5524
  -2.500   0.0535   0.00804   0.00258  -0.0759   0.7162   0.5656
  -2.250   0.0826   0.00805   0.00254  -0.0761   0.7101   0.5754
  -2.000   0.1115   0.00809   0.00255  -0.0763   0.7043   0.5863
  -1.750   0.1404   0.00807   0.00255  -0.0764   0.6977   0.5958
  -1.500   0.1694   0.00811   0.00252  -0.0766   0.6919   0.6037
  -1.250   0.1984   0.00811   0.00252  -0.0768   0.6860   0.6104
  -1.000   0.2274   0.00811   0.00252  -0.0770   0.6797   0.6175
  -0.750   0.2564   0.00816   0.00250  -0.0772   0.6741   0.6244
  -0.500   0.2852   0.00815   0.00253  -0.0774   0.6679   0.6309
  -0.250   0.3144   0.00819   0.00253  -0.0776   0.6620   0.6380
   0.000   0.3431   0.00822   0.00255  -0.0778   0.6566   0.6441
   0.250   0.3720   0.00823   0.00259  -0.0780   0.6503   0.6507
   0.500   0.4011   0.00827   0.00260  -0.0782   0.6445   0.6573
   0.750   0.4297   0.00831   0.00265  -0.0784   0.6390   0.6632
   1.000   0.4587   0.00834   0.00270  -0.0786   0.6327   0.6702
   1.250   0.4874   0.00838   0.00273  -0.0787   0.6270   0.6765
   1.500   0.5160   0.00843   0.00280  -0.0789   0.6211   0.6828
   1.750   0.5449   0.00848   0.00285  -0.0791   0.6147   0.6899
   2.000   0.5732   0.00855   0.00292  -0.0792   0.6090   0.6961
   2.250   0.6018   0.00859   0.00301  -0.0793   0.6027   0.7031
   2.500   0.6304   0.00864   0.00308  -0.0795   0.5965   0.7097
   2.750   0.6586   0.00873   0.00318  -0.0795   0.5907   0.7162
   3.000   0.6873   0.00878   0.00327  -0.0797   0.5838   0.7238
   3.250   0.7150   0.00886   0.00335  -0.0797   0.5768   0.7302
   3.500   0.7430   0.00889   0.00345  -0.0797   0.5678   0.7375
   3.750   0.7706   0.00898   0.00353  -0.0797   0.5594   0.7443
   4.000   0.7983   0.00902   0.00364  -0.0796   0.5508   0.7516
   4.250   0.8258   0.00912   0.00374  -0.0796   0.5416   0.7591
   4.500   0.8526   0.00918   0.00385  -0.0794   0.5300   0.7661
   4.750   0.8798   0.00927   0.00395  -0.0793   0.5163   0.7742
   5.000   0.9057   0.00936   0.00408  -0.0789   0.5011   0.7814
   5.250   0.9324   0.00950   0.00422  -0.0788   0.4867   0.7897
   5.500   0.9582   0.00963   0.00439  -0.0784   0.4730   0.7971
   5.750   0.9843   0.00981   0.00458  -0.0781   0.4575   0.8058
   6.000   1.0089   0.01000   0.00478  -0.0776   0.4385   0.8136
   6.250   1.0330   0.01029   0.00501  -0.0770   0.4132   0.8227
   6.500   1.0546   0.01066   0.00531  -0.0760   0.3793   0.8311
   6.750   1.0739   0.01125   0.00571  -0.0748   0.3334   0.8410
   7.000   1.0892   0.01205   0.00627  -0.0730   0.2796   0.8508
   7.250   1.1023   0.01297   0.00692  -0.0708   0.2239   0.8615
   7.500   1.1139   0.01393   0.00763  -0.0685   0.1737   0.8737
   7.750   1.1250   0.01479   0.00831  -0.0661   0.1343   0.8874
   8.000   1.1333   0.01562   0.00898  -0.0630   0.1020   0.9038
   8.250   1.1375   0.01628   0.00958  -0.0591   0.0796   0.9283
   8.500   1.1435   0.01693   0.01019  -0.0558   0.0645   1.0000
   8.750   1.1562   0.01780   0.01099  -0.0541   0.0548   1.0000
   9.000   1.1708   0.01854   0.01174  -0.0527   0.0497   1.0000
   9.250   1.1818   0.01948   0.01266  -0.0509   0.0456   1.0000
   9.500   1.1957   0.02027   0.01349  -0.0495   0.0429   1.0000
   9.750   1.2056   0.02132   0.01455  -0.0478   0.0403   1.0000
  10.000   1.2147   0.02248   0.01574  -0.0461   0.0382   1.0000
  10.250   1.2273   0.02345   0.01676  -0.0449   0.0363   1.0000
  10.500   1.2374   0.02462   0.01797  -0.0436   0.0345   1.0000
  10.750   1.2391   0.02645   0.01981  -0.0418   0.0326   1.0000
  11.000   1.2523   0.02751   0.02095  -0.0409   0.0314   1.0000
  11.250   1.2625   0.02882   0.02231  -0.0399   0.0298   1.0000
  11.500   1.2704   0.03035   0.02388  -0.0389   0.0287   1.0000
  11.750   1.2701   0.03262   0.02617  -0.0375   0.0274   1.0000
  12.000   1.2778   0.03429   0.02792  -0.0367   0.0266   1.0000
  12.250   1.2867   0.03591   0.02962  -0.0360   0.0257   1.0000
  12.500   1.2940   0.03770   0.03147  -0.0353   0.0248   1.0000
  12.750   1.3005   0.03960   0.03342  -0.0347   0.0241   1.0000
  13.000   1.3047   0.04176   0.03561  -0.0342   0.0234   1.0000
  13.250   1.3020   0.04464   0.03854  -0.0334   0.0226   1.0000
  13.500   1.3105   0.04652   0.04051  -0.0331   0.0222   1.0000
  13.750   1.3174   0.04860   0.04268  -0.0329   0.0217   1.0000
  14.000   1.3239   0.05076   0.04493  -0.0327   0.0210   1.0000
  14.250   1.3297   0.05305   0.04730  -0.0327   0.0205   1.0000
  14.500   1.3357   0.05537   0.04968  -0.0327   0.0201   1.0000
  14.750   1.3406   0.05786   0.05221  -0.0329   0.0196   1.0000
  15.000   1.3446   0.06048   0.05488  -0.0330   0.0193   1.0000
  15.250   1.3475   0.06325   0.05769  -0.0330   0.0189   1.0000
  15.500   1.3510   0.06603   0.06056  -0.0330   0.0185   1.0000
  15.750   1.3554   0.06881   0.06346  -0.0334   0.0183   1.0000
  16.000   1.3588   0.07178   0.06656  -0.0340   0.0181   1.0000
  16.250   1.3619   0.07485   0.06975  -0.0346   0.0178   1.0000
  16.500   1.3642   0.07809   0.07310  -0.0354   0.0175   1.0000
  16.750   1.3660   0.08144   0.07657  -0.0362   0.0173   1.0000
  17.000   1.3668   0.08499   0.08023  -0.0373   0.0170   1.0000
  17.250   1.3669   0.08871   0.08407  -0.0385   0.0168   1.0000
  17.500   1.3663   0.09262   0.08810  -0.0399   0.0166   1.0000
  17.750   1.3653   0.09665   0.09224  -0.0414   0.0164   1.0000
  18.000   1.3637   0.10083   0.09652  -0.0431   0.0163   1.0000
  18.250   1.3615   0.10515   0.10095  -0.0450   0.0161   1.0000
  18.500   1.3595   0.10951   0.10541  -0.0470   0.0159   1.0000
  18.750   1.3572   0.11393   0.10991  -0.0491   0.0158   1.0000
  19.000   1.3540   0.11852   0.11460  -0.0514   0.0156   1.0000
  19.250   1.3489   0.12347   0.11966  -0.0538   0.0154   1.0000
<< Back to NACA 63-415 AIRFOIL (n63415-il)

Polar data table (+)

Polar graphs


<< Back to NACA 63-415 AIRFOIL (n63415-il)