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NACA M19 AIRFOIL (m19-il) Xfoil prediction polar at RE=200,000 Ncrit=9


Details Polar file
Airfoil: NACA M19 AIRFOIL (m19-il)
Reynolds number: 200,000
Max Cl/Cd: 77.39 at α=6°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-m19-il-200000.txt
Download as CSV file: xf-m19-il-200000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NACA M19 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
  ------ -------- --------- --------- -------- -------- --------
  -8.250  -0.2723   0.09572   0.09275  -0.0055   0.8434   0.0214
  -8.000  -0.2682   0.09253   0.08952  -0.0058   0.8320   0.0219
  -7.750  -0.3600   0.10342   0.10047  -0.0015   0.8653   0.0206
  -7.500  -0.3536   0.09932   0.09628  -0.0005   0.8511   0.0210
  -7.250  -0.3472   0.09613   0.09303  -0.0011   0.8380   0.0215
  -7.000  -0.3384   0.09308   0.08991  -0.0026   0.8265   0.0222
  -6.750  -0.3269   0.08999   0.08676  -0.0050   0.8153   0.0232
  -6.500  -0.3135   0.08681   0.08353  -0.0079   0.8050   0.0241
  -6.250  -0.2983   0.08363   0.08027  -0.0110   0.7959   0.0249
  -6.000  -0.2799   0.08051   0.07708  -0.0147   0.7869   0.0260
  -5.750  -0.2514   0.07829   0.07475  -0.0209   0.7777   0.0271
  -4.750  -0.1564   0.06394   0.05992  -0.0353   0.7472   0.0284
  -4.500  -0.1467   0.05997   0.05593  -0.0343   0.7407   0.0296
  -4.250  -0.1257   0.05697   0.05285  -0.0358   0.7328   0.0313
  -4.000  -0.1001   0.05409   0.04982  -0.0379   0.7265   0.0331
  -3.750  -0.0701   0.05136   0.04695  -0.0405   0.7188   0.0356
  -3.500  -0.0196   0.05125   0.04639  -0.0447   0.7124   0.0381
  -3.250  -0.0025   0.04572   0.04082  -0.0458   0.7055   0.0393
  -3.000   0.0136   0.04285   0.03788  -0.0456   0.6998   0.0416
  -2.750   0.0407   0.04060   0.03552  -0.0468   0.6921   0.0451
  -2.500   0.0876   0.04108   0.03543  -0.0483   0.6861   0.0506
  -2.250   0.1047   0.03613   0.03053  -0.0491   0.6792   0.0525
  -2.000   0.1272   0.03407   0.02837  -0.0493   0.6730   0.0556
  -1.750   0.1654   0.03339   0.02722  -0.0500   0.6663   0.0645
  -1.500   0.1862   0.03058   0.02442  -0.0504   0.6599   0.0668
  -1.250   0.2125   0.02908   0.02279  -0.0506   0.6535   0.0718
  -1.000   0.2443   0.02772   0.02110  -0.0508   0.6469   0.0793
  -0.750   0.2693   0.02620   0.01948  -0.0510   0.6411   0.0841
  -0.500   0.2988   0.02499   0.01805  -0.0512   0.6340   0.0941
  -0.250   0.3273   0.02419   0.01696  -0.0511   0.6287   0.1065
   0.000   0.3538   0.02286   0.01563  -0.0514   0.6213   0.1126
   0.250   0.3817   0.02187   0.01444  -0.0513   0.6156   0.1236
   0.500   0.4097   0.02101   0.01348  -0.0515   0.6088   0.1373
   0.750   0.4372   0.02020   0.01253  -0.0515   0.6028   0.1536
   1.000   0.4644   0.01941   0.01166  -0.0516   0.5966   0.1790
   1.250   0.5026   0.01740   0.00887  -0.0500   0.5908   0.0737
   1.500   0.5305   0.01680   0.00805  -0.0497   0.5858   0.0741
   1.750   0.5589   0.01613   0.00735  -0.0496   0.5786   0.0718
   2.000   0.5866   0.01556   0.00663  -0.0493   0.5733   0.0685
   2.250   0.6141   0.01517   0.00623  -0.0491   0.5670   0.0670
   2.500   0.6410   0.01486   0.00589  -0.0489   0.5609   0.0668
   2.750   0.6679   0.01465   0.00570  -0.0487   0.5552   0.0680
   3.000   0.6950   0.01450   0.00560  -0.0486   0.5486   0.0713
   3.250   0.7218   0.01445   0.00548  -0.0483   0.5439   0.0776
   3.500   0.7692   0.01280   0.00561  -0.0528   0.5365   1.0000
   3.750   0.7955   0.01296   0.00562  -0.0524   0.5313   1.0000
   4.000   0.8219   0.01316   0.00581  -0.0523   0.5252   1.0000
   4.250   0.8483   0.01334   0.00599  -0.0521   0.5193   1.0000
   4.500   0.8746   0.01355   0.00614  -0.0519   0.5143   1.0000
   4.750   0.9002   0.01338   0.00591  -0.0513   0.5009   1.0000
   5.000   0.9258   0.01316   0.00569  -0.0508   0.4825   1.0000
   5.250   0.9516   0.01319   0.00574  -0.0505   0.4712   1.0000
   5.500   0.9773   0.01313   0.00573  -0.0501   0.4531   1.0000
   5.750   1.0023   0.01307   0.00560  -0.0496   0.4207   1.0000
   6.000   1.0270   0.01327   0.00573  -0.0492   0.3819   1.0000
   6.250   1.0478   0.01423   0.00617  -0.0487   0.2824   1.0000
   6.500   1.0513   0.01825   0.00875  -0.0476   0.0377   1.0000
   6.750   1.0731   0.01908   0.00970  -0.0469   0.0322   1.0000
   7.000   1.0941   0.01995   0.01076  -0.0462   0.0303   1.0000
   7.250   1.1131   0.02101   0.01202  -0.0453   0.0291   1.0000
   7.500   1.1295   0.02228   0.01349  -0.0442   0.0285   1.0000
   7.750   1.1419   0.02383   0.01517  -0.0429   0.0271   1.0000
   8.000   1.1485   0.02575   0.01718  -0.0412   0.0255   1.0000
   8.250   1.1526   0.02773   0.01927  -0.0392   0.0249   1.0000
   8.500   1.1556   0.02957   0.02118  -0.0370   0.0248   1.0000
   8.750   1.1610   0.03141   0.02306  -0.0350   0.0249   1.0000
   9.000   1.1717   0.03299   0.02469  -0.0331   0.0253   1.0000
   9.250   1.1884   0.03432   0.02610  -0.0311   0.0263   1.0000
   9.500   1.2261   0.03592   0.02781  -0.0292   0.0299   1.0000
   9.750   1.2930   0.03928   0.03135  -0.0300   0.0416   1.0000
  16.250   0.7879   0.16938   0.16618  -0.0438   0.0783   1.0000
  16.500   0.8147   0.16776   0.16460  -0.0386   0.0673   1.0000
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