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

NACA M27 AIRFOIL (m27-il) Xfoil prediction polar at RE=200,000 Ncrit=5


Details Polar file
Airfoil: NACA M27 AIRFOIL (m27-il)
Reynolds number: 200,000
Max Cl/Cd: 70.12 at α=7.5°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-m27-il-200000-n5.txt
Download as CSV file: xf-m27-il-200000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NACA M27 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 =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -7.750  -0.1484   0.09594   0.09124  -0.0217   0.5886   0.0275
  -7.500  -0.1414   0.09310   0.08836  -0.0239   0.5860   0.0283
  -7.250  -0.1362   0.09083   0.08610  -0.0275   0.5828   0.0295
  -7.000  -0.1246   0.08869   0.08391  -0.0335   0.5795   0.0299
  -6.750  -0.1103   0.08601   0.08116  -0.0377   0.5763   0.0300
  -6.500  -0.0978   0.08220   0.07731  -0.0386   0.5732   0.0301
  -6.250  -0.0870   0.07851   0.07357  -0.0379   0.5704   0.0304
  -6.000  -0.0734   0.07553   0.07059  -0.0387   0.5670   0.0309
  -5.750  -0.0578   0.07281   0.06784  -0.0401   0.5637   0.0315
  -5.500  -0.0407   0.07023   0.06521  -0.0418   0.5604   0.0325
  -5.250  -0.0214   0.06768   0.06256  -0.0439   0.5574   0.0343
  -5.000   0.0085   0.06566   0.06032  -0.0484   0.5547   0.0361
  -4.750   0.0371   0.06344   0.05791  -0.0518   0.5515   0.0363
  -4.500   0.0584   0.06036   0.05475  -0.0532   0.5481   0.0365
  -4.250   0.0718   0.05699   0.05139  -0.0528   0.5448   0.0369
  -4.000   0.0889   0.05447   0.04882  -0.0529   0.5419   0.0376
  -3.750   0.1093   0.05239   0.04664  -0.0534   0.5392   0.0389
  -3.500   0.1329   0.05038   0.04452  -0.0543   0.5364   0.0407
  -3.250   0.1652   0.04873   0.04268  -0.0562   0.5329   0.0434
  -3.000   0.1997   0.04724   0.04090  -0.0577   0.5296   0.0440
  -2.750   0.2206   0.04440   0.03796  -0.0579   0.5266   0.0443
  -2.500   0.2391   0.04204   0.03555  -0.0578   0.5239   0.0449
  -2.250   0.2610   0.04026   0.03366  -0.0578   0.5215   0.0458
  -2.000   0.2853   0.03877   0.03212  -0.0581   0.5182   0.0478
  -1.750   0.3242   0.03828   0.03126  -0.0587   0.5149   0.0524
  -1.500   0.3499   0.03628   0.02908  -0.0587   0.5119   0.0527
  -1.250   0.3707   0.03404   0.02679  -0.0587   0.5090   0.0534
  -1.000   0.3944   0.03253   0.02516  -0.0586   0.5065   0.0541
  -0.750   0.4244   0.03010   0.02240  -0.0580   0.5038   0.0435
  -0.500   0.4489   0.02874   0.02099  -0.0580   0.5008   0.0418
  -0.250   0.4763   0.02728   0.01935  -0.0578   0.4978   0.0406
   0.000   0.5043   0.02597   0.01783  -0.0576   0.4949   0.0415
   0.250   0.5321   0.02459   0.01620  -0.0573   0.4923   0.0405
   0.500   0.5608   0.02293   0.01420  -0.0569   0.4899   0.0390
   0.750   0.5904   0.02109   0.01204  -0.0566   0.4869   0.0381
   1.000   0.6195   0.01997   0.01069  -0.0567   0.4839   0.0381
   1.250   0.6482   0.01926   0.00981  -0.0569   0.4810   0.0385
   1.500   0.6767   0.01873   0.00914  -0.0570   0.4782   0.0391
   1.750   0.7046   0.01842   0.00872  -0.0571   0.4757   0.0410
   2.000   0.7325   0.01808   0.00826  -0.0572   0.4731   0.0425
   2.250   0.7596   0.01778   0.00795  -0.0573   0.4698   0.0430
   2.500   0.7861   0.01757   0.00772  -0.0572   0.4668   0.0437
   2.750   0.8119   0.01738   0.00753  -0.0570   0.4641   0.0448
   3.000   0.8379   0.01728   0.00744  -0.0568   0.4615   0.0465
   3.250   0.8639   0.01724   0.00735  -0.0567   0.4592   0.0489
   3.500   0.8896   0.01726   0.00740  -0.0566   0.4564   0.0532
   3.750   0.9145   0.01731   0.00752  -0.0563   0.4532   0.0607
   4.000   0.9397   0.01735   0.00761  -0.0561   0.4502   0.0754
   4.250   0.9654   0.01743   0.00771  -0.0560   0.4476   0.1013
   4.500   0.9912   0.01751   0.00783  -0.0560   0.4452   0.1337
   4.750   1.0171   0.01757   0.00796  -0.0559   0.4430   0.2059
   5.250   1.1847   0.01737   0.00915  -0.0810   0.4346   1.0000
   5.500   1.2081   0.01759   0.00936  -0.0806   0.4321   1.0000
   5.750   1.2314   0.01781   0.00954  -0.0801   0.4298   1.0000
   6.000   1.2549   0.01803   0.00971  -0.0797   0.4277   1.0000
   6.250   1.2770   0.01833   0.01006  -0.0792   0.4249   1.0000
   6.500   1.2986   0.01865   0.01043  -0.0786   0.4218   1.0000
   6.750   1.3202   0.01894   0.01076  -0.0779   0.4189   1.0000
   7.000   1.3418   0.01922   0.01108  -0.0773   0.4164   1.0000
   7.250   1.3635   0.01949   0.01134  -0.0766   0.4142   1.0000
   7.500   1.3856   0.01976   0.01159  -0.0760   0.4122   1.0000
   7.750   1.4050   0.02016   0.01209  -0.0751   0.4095   1.0000
   8.000   1.4228   0.02055   0.01258  -0.0740   0.4055   1.0000
   8.250   1.4404   0.02081   0.01286  -0.0728   0.4003   1.0000
   8.500   1.4582   0.02104   0.01307  -0.0715   0.3950   1.0000
   8.750   1.4709   0.02152   0.01369  -0.0697   0.3893   1.0000
   9.000   1.4853   0.02190   0.01409  -0.0680   0.3847   1.0000
   9.250   1.4979   0.02230   0.01450  -0.0661   0.3795   1.0000
   9.500   1.5032   0.02295   0.01526  -0.0632   0.3730   1.0000
   9.750   1.5065   0.02354   0.01585  -0.0600   0.3675   1.0000
  10.000   1.5034   0.02443   0.01682  -0.0561   0.3627   1.0000
  10.250   1.5010   0.02556   0.01804  -0.0530   0.3566   1.0000
  10.500   1.5025   0.02670   0.01916  -0.0506   0.3501   1.0000
  10.750   1.5010   0.02833   0.02091  -0.0486   0.3427   1.0000
  11.000   1.5024   0.02985   0.02243  -0.0469   0.3357   1.0000
  11.250   1.5008   0.03187   0.02457  -0.0454   0.3276   1.0000
  11.500   1.4959   0.03421   0.02693  -0.0439   0.3170   1.0000
  11.750   1.4889   0.03685   0.02957  -0.0425   0.3059   1.0000
  12.000   1.4801   0.03985   0.03262  -0.0413   0.2934   1.0000
  12.250   1.4685   0.04317   0.03595  -0.0401   0.2795   1.0000
  12.500   1.4545   0.04687   0.03964  -0.0389   0.2644   1.0000
  12.750   1.4389   0.05095   0.04369  -0.0380   0.2485   1.0000
  13.000   1.4216   0.05540   0.04809  -0.0372   0.2298   1.0000
  13.250   1.3995   0.06052   0.05314  -0.0365   0.2090   1.0000
  13.500   1.3719   0.06650   0.05901  -0.0360   0.1856   1.0000
  13.750   1.3450   0.07258   0.06499  -0.0357   0.1632   1.0000
  14.000   1.3180   0.07887   0.07118  -0.0356   0.1396   1.0000
  14.250   1.2914   0.08527   0.07748  -0.0357   0.1137   1.0000
  14.500   1.2658   0.09172   0.08380  -0.0359   0.0847   1.0000
<< Back to NACA M27 AIRFOIL (m27-il)

Polar data table (+)

Polar graphs


<< Back to NACA M27 AIRFOIL (m27-il)