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

NACA M25 AIRFOIL (m25-il) Xfoil prediction polar at RE=200,000 Ncrit=9


Details Polar file
Airfoil: NACA M25 AIRFOIL (m25-il)
Reynolds number: 200,000
Max Cl/Cd: 78.36 at α=7.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-m25-il-200000.txt
Download as CSV file: xf-m25-il-200000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NACA M25 AIRFOIL                                
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.200 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.500  -0.3409   0.11989   0.11620   0.0197   0.6957   0.0179
  -8.250  -0.3328   0.11783   0.11416   0.0169   0.6903   0.0179
  -8.000  -0.3258   0.11584   0.11217   0.0139   0.6854   0.0180
  -7.750  -0.3179   0.11395   0.11023   0.0107   0.6815   0.0180
  -7.250  -0.2916   0.10817   0.10441   0.0037   0.6716   0.0181
  -7.000  -0.2865   0.10115   0.09736   0.0069   0.6678   0.0183
  -6.750  -0.2765   0.09712   0.09331   0.0061   0.6638   0.0186
  -6.500  -0.2638   0.09370   0.08989   0.0040   0.6589   0.0189
  -6.250  -0.2497   0.09055   0.08669   0.0017   0.6548   0.0193
  -6.000  -0.2342   0.08751   0.08356  -0.0008   0.6514   0.0198
  -5.750  -0.2157   0.08437   0.08043  -0.0039   0.6463   0.0204
  -5.500  -0.1955   0.08140   0.07740  -0.0071   0.6419   0.0216
  -5.250  -0.1663   0.07948   0.07535  -0.0118   0.6382   0.0228
  -5.000  -0.1307   0.07831   0.07404  -0.0180   0.6340   0.0231
  -4.750  -0.0978   0.07651   0.07213  -0.0230   0.6292   0.0232
  -4.250  -0.0590   0.06816   0.06360  -0.0260   0.6224   0.0236
  -4.000  -0.0462   0.06356   0.05904  -0.0262   0.6180   0.0240
  -3.750  -0.0257   0.06031   0.05575  -0.0276   0.6138   0.0247
  -3.500  -0.0013   0.05755   0.05287  -0.0294   0.6102   0.0256
  -3.250   0.0257   0.05509   0.05025  -0.0313   0.6069   0.0275
  -3.000   0.0658   0.05394   0.04896  -0.0347   0.6018   0.0299
  -2.750   0.1108   0.05383   0.04856  -0.0380   0.5977   0.0305
  -2.500   0.1345   0.04985   0.04444  -0.0391   0.5946   0.0310
  -2.250   0.1494   0.04572   0.04031  -0.0393   0.5911   0.0318
  -2.000   0.1731   0.04324   0.03778  -0.0402   0.5865   0.0335
  -1.750   0.2014   0.04135   0.03575  -0.0412   0.5827   0.0364
  -1.500   0.2439   0.04161   0.03564  -0.0426   0.5795   0.0401
  -1.250   0.2747   0.03923   0.03310  -0.0436   0.5755   0.0411
  -1.000   0.2941   0.03607   0.02996  -0.0441   0.5711   0.0426
  -0.750   0.3194   0.03436   0.02812  -0.0445   0.5675   0.0459
  -0.500   0.3595   0.03517   0.02852  -0.0447   0.5645   0.0523
  -0.250   0.3852   0.03233   0.02563  -0.0455   0.5602   0.0537
   0.000   0.4087   0.03036   0.02365  -0.0460   0.5559   0.0561
   0.250   0.4366   0.02923   0.02235  -0.0462   0.5524   0.0610
   0.500   0.4700   0.02870   0.02145  -0.0460   0.5496   0.0666
   0.750   0.4946   0.02713   0.01997  -0.0468   0.5450   0.0710
   1.000   0.5285   0.02717   0.01971  -0.0467   0.5409   0.0791
   1.250   0.5526   0.02533   0.01785  -0.0470   0.5376   0.0821
   1.500   0.5837   0.02541   0.01759  -0.0467   0.5345   0.0922
   1.750   0.6094   0.02385   0.01616  -0.0474   0.5297   0.0962
   2.000   0.6384   0.02330   0.01546  -0.0475   0.5258   0.1074
   2.250   0.6663   0.02273   0.01473  -0.0474   0.5228   0.1205
   2.500   0.6933   0.02205   0.01396  -0.0475   0.5196   0.1355
   2.750   0.7209   0.02164   0.01359  -0.0480   0.5146   0.1624
   3.000   0.7479   0.02104   0.01296  -0.0481   0.5108   0.1903
   3.250   0.7748   0.02077   0.01258  -0.0481   0.5078   0.2297
   3.500   0.8019   0.01993   0.01179  -0.0483   0.5044   0.2525
   3.750   0.8397   0.01918   0.01051  -0.0470   0.5004   0.0713
   4.000   0.8674   0.01881   0.01002  -0.0468   0.4967   0.0667
   4.250   0.8941   0.01849   0.00964  -0.0465   0.4937   0.0666
   4.500   0.9204   0.01834   0.00952  -0.0464   0.4904   0.0687
   4.750   0.9470   0.01845   0.00976  -0.0466   0.4855   0.0765
   5.000   0.9729   0.01837   0.00968  -0.0463   0.4820   0.0796
   5.250   0.9998   0.01835   0.00964  -0.0462   0.4792   0.0875
   5.500   1.0268   0.01842   0.00986  -0.0463   0.4762   0.1985
   5.750   1.0966   0.01755   0.01023  -0.0561   0.4685   1.0000
   6.000   1.1219   0.01696   0.00953  -0.0553   0.4580   1.0000
   6.250   1.1474   0.01655   0.00901  -0.0546   0.4494   1.0000
   6.500   1.1727   0.01626   0.00878  -0.0543   0.4373   1.0000
   6.750   1.1979   0.01591   0.00842  -0.0539   0.4221   1.0000
   7.000   1.2229   0.01567   0.00822  -0.0537   0.3975   1.0000
   7.250   1.2460   0.01590   0.00826  -0.0536   0.3499   1.0000
   7.500   1.2598   0.01793   0.00958  -0.0536   0.2355   1.0000
   7.750   1.2441   0.02322   0.01380  -0.0525   0.0449   1.0000
   8.000   1.2542   0.02474   0.01534  -0.0514   0.0360   1.0000
   8.250   1.2632   0.02631   0.01709  -0.0507   0.0338   1.0000
   8.500   1.2634   0.02844   0.01938  -0.0495   0.0319   1.0000
   8.750   1.2598   0.03092   0.02201  -0.0481   0.0302   1.0000
   9.000   1.2560   0.03376   0.02500  -0.0473   0.0289   1.0000
   9.250   1.2516   0.03683   0.02824  -0.0467   0.0283   1.0000
   9.500   1.2452   0.04022   0.03178  -0.0462   0.0278   1.0000
   9.750   1.2370   0.04387   0.03557  -0.0457   0.0275   1.0000
  10.000   1.2280   0.04764   0.03951  -0.0454   0.0273   1.0000
  10.250   1.2185   0.05149   0.04347  -0.0450   0.0271   1.0000
  10.500   1.2097   0.05523   0.04730  -0.0445   0.0269   1.0000
  10.750   1.2035   0.05866   0.05080  -0.0440   0.0268   1.0000
  11.000   1.2008   0.06162   0.05381  -0.0432   0.0267   1.0000
  11.250   1.2024   0.06393   0.05614  -0.0420   0.0267   1.0000
  11.500   1.2089   0.06519   0.05737  -0.0398   0.0265   1.0000
  11.750   1.2176   0.06697   0.05923  -0.0391   0.0258   1.0000
  12.000   1.2303   0.06776   0.06004  -0.0371   0.0253   1.0000
  12.250   1.2843   0.06311   0.05524  -0.0292   0.0269   1.0000
  12.500   1.2924   0.06497   0.05722  -0.0290   0.0278   1.0000
  12.750   1.1305   0.06581   0.05865  -0.0164   0.0265   1.0000
  13.000   1.1404   0.06761   0.06056  -0.0156   0.0272   1.0000
<< Back to NACA M25 AIRFOIL (m25-il)

Polar data table (+)

Polar graphs


<< Back to NACA M25 AIRFOIL (m25-il)