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NACA 63-210 AIRFOIL (n63210-il) Xfoil prediction polar at RE=100,000 Ncrit=9


Details Polar file
Airfoil: NACA 63-210 AIRFOIL (n63210-il)
Reynolds number: 100,000
Max Cl/Cd: 46.66 at α=4.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-n63210-il-100000.txt
Download as CSV file: xf-n63210-il-100000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NACA 63-210 AIRFOIL                             
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.100 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.500  -0.4652   0.09208   0.08751  -0.0194   1.0000   0.1467
  -9.250  -0.4470   0.08783   0.08324  -0.0176   1.0000   0.1524
  -9.000  -0.5520   0.09036   0.08557  -0.0198   1.0000   0.1458
  -8.750  -0.5424   0.08670   0.08193  -0.0188   1.0000   0.1531
  -8.500  -0.5625   0.08227   0.07761  -0.0239   1.0000   0.1602
  -8.250  -0.5491   0.07902   0.07437  -0.0218   1.0000   0.1684
  -8.000  -0.5866   0.07352   0.06896  -0.0300   1.0000   0.1750
  -7.750  -0.5673   0.07069   0.06618  -0.0267   1.0000   0.1855
  -7.500  -0.5701   0.06687   0.06238  -0.0277   1.0000   0.1972
  -7.250  -0.5686   0.06382   0.05935  -0.0275   1.0000   0.2132
  -6.750  -0.5938   0.04111   0.03423  -0.0386   1.0000   0.0922
  -6.500  -0.5765   0.03760   0.03049  -0.0374   1.0000   0.0888
  -6.250  -0.5601   0.03376   0.02632  -0.0361   1.0000   0.0837
  -6.000  -0.5403   0.03175   0.02338  -0.0339   1.0000   0.0775
  -5.750  -0.5213   0.02941   0.02074  -0.0326   1.0000   0.0772
  -5.500  -0.5029   0.02681   0.01794  -0.0314   1.0000   0.0786
  -5.250  -0.4830   0.02512   0.01604  -0.0299   1.0000   0.0787
  -5.000  -0.4629   0.02364   0.01439  -0.0284   1.0000   0.0786
  -4.750  -0.4430   0.02221   0.01287  -0.0270   1.0000   0.0792
  -4.500  -0.4236   0.02099   0.01164  -0.0256   1.0000   0.0804
  -4.250  -0.4045   0.01999   0.01067  -0.0242   1.0000   0.0824
  -4.000  -0.3857   0.01914   0.00984  -0.0228   1.0000   0.0853
  -3.750  -0.3668   0.01844   0.00911  -0.0215   1.0000   0.0893
  -3.500  -0.3477   0.01764   0.00838  -0.0205   1.0000   0.0975
  -3.250  -0.3274   0.01695   0.00772  -0.0198   1.0000   0.1106
  -3.000  -0.3088   0.01383   0.00690  -0.0197   1.0000   0.5082
  -2.750  -0.2963   0.01387   0.00735  -0.0160   1.0000   0.6539
  -2.500  -0.2821   0.01414   0.00767  -0.0132   1.0000   0.7036
  -2.250  -0.2664   0.01466   0.00830  -0.0099   0.9964   0.7606
  -2.000  -0.2472   0.01550   0.00926  -0.0058   0.9857   0.8201
  -1.750  -0.2293   0.01606   0.00984  -0.0017   0.9750   0.8626
  -1.500  -0.2007   0.01632   0.01002  -0.0010   0.9658   0.8894
  -1.250  -0.1613   0.01645   0.00999  -0.0033   0.9576   0.9067
  -1.000  -0.1228   0.01649   0.00993  -0.0056   0.9490   0.9229
  -0.750  -0.0702   0.01659   0.00989  -0.0107   0.9430   0.9366
  -0.500  -0.0212   0.01662   0.00984  -0.0153   0.9352   0.9498
  -0.250   0.0408   0.01666   0.00979  -0.0224   0.9302   0.9598
   0.000   0.1041   0.01665   0.00972  -0.0298   0.9250   0.9687
   0.250   0.1702   0.01654   0.00959  -0.0378   0.9193   0.9756
   0.500   0.2354   0.01635   0.00940  -0.0455   0.9146   0.9831
   0.750   0.2945   0.01613   0.00922  -0.0523   0.9067   0.9921
   1.000   0.3560   0.01579   0.00895  -0.0593   0.9005   1.0000
   1.250   0.3799   0.01574   0.00893  -0.0594   0.8877   1.0000
   1.500   0.4011   0.01575   0.00897  -0.0589   0.8749   1.0000
   1.750   0.4203   0.01583   0.00908  -0.0578   0.8621   1.0000
   2.000   0.4380   0.01596   0.00928  -0.0563   0.8495   1.0000
   2.250   0.4553   0.01614   0.00949  -0.0547   0.8370   1.0000
   2.500   0.4735   0.01631   0.00970  -0.0530   0.8246   1.0000
   2.750   0.4944   0.01642   0.00986  -0.0514   0.8121   1.0000
   3.000   0.5170   0.01648   0.01001  -0.0500   0.7991   1.0000
   3.250   0.5384   0.01648   0.01008  -0.0482   0.7820   1.0000
   3.500   0.5612   0.01586   0.00946  -0.0451   0.7572   1.0000
   3.750   0.5825   0.01502   0.00857  -0.0413   0.7235   1.0000
   4.000   0.6026   0.01425   0.00772  -0.0376   0.6785   1.0000
   4.250   0.6235   0.01388   0.00725  -0.0348   0.6274   1.0000
   4.500   0.6448   0.01382   0.00711  -0.0327   0.5662   1.0000
   4.750   0.6472   0.01560   0.00717  -0.0278   0.2419   1.0000
   5.000   0.6553   0.01846   0.00888  -0.0255   0.1020   1.0000
   5.250   0.6747   0.01972   0.01012  -0.0242   0.0900   1.0000
   5.500   0.6955   0.02089   0.01125  -0.0231   0.0834   1.0000
   5.750   0.7174   0.02232   0.01258  -0.0222   0.0786   1.0000
   6.000   0.7424   0.02350   0.01381  -0.0216   0.0739   1.0000
   6.250   0.7689   0.02499   0.01530  -0.0212   0.0716   1.0000
   6.500   0.7966   0.02674   0.01710  -0.0210   0.0700   1.0000
   6.750   0.8248   0.02877   0.01923  -0.0208   0.0692   1.0000
   7.000   0.8523   0.03100   0.02166  -0.0204   0.0691   1.0000
   7.250   0.8784   0.03336   0.02438  -0.0198   0.0699   1.0000
   7.500   0.9011   0.03617   0.02773  -0.0186   0.0712   1.0000
   7.750   0.9219   0.03891   0.03081  -0.0178   0.0703   1.0000
   8.000   0.9370   0.04287   0.03549  -0.0161   0.0732   1.0000
   8.250   0.9513   0.04767   0.04072  -0.0148   0.0777   1.0000
   8.500   0.9611   0.05480   0.04849  -0.0132   0.0925   1.0000
  10.500   0.8305   0.11135   0.10695  -0.0275   0.1632   1.0000
  10.750   0.7691   0.11917   0.11455  -0.0403   0.1562   1.0000
  11.000   0.6480   0.12249   0.11821  -0.0356   0.1605   1.0000
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