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

NACA M20 AIRFOIL (m20-il) Xfoil prediction polar at RE=200,000 Ncrit=9


Details Polar file
Airfoil: NACA M20 AIRFOIL (m20-il)
Reynolds number: 200,000
Max Cl/Cd: 78.64 at α=7.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-m20-il-200000.txt
Download as CSV file: xf-m20-il-200000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NACA M20 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
  ------ -------- --------- --------- -------- -------- --------
  -9.500  -0.2543   0.10976   0.10674  -0.0101   0.8576   0.0271
  -9.250  -0.2514   0.10683   0.10376  -0.0106   0.8463   0.0278
  -9.000  -0.3503   0.11388   0.11090  -0.0013   0.9023   0.0260
  -8.750  -0.3473   0.11130   0.10825  -0.0014   0.8800   0.0266
  -8.500  -0.3429   0.10856   0.10544  -0.0021   0.8629   0.0272
  -8.250  -0.3379   0.10578   0.10259  -0.0031   0.8488   0.0279
  -8.000  -0.3326   0.10311   0.09986  -0.0046   0.8365   0.0288
  -7.750  -0.3274   0.10132   0.09803  -0.0087   0.8240   0.0297
  -7.500  -0.3219   0.10011   0.09677  -0.0158   0.8123   0.0300
  -7.250  -0.3077   0.09783   0.09439  -0.0221   0.8028   0.0302
  -7.000  -0.2980   0.09302   0.08952  -0.0240   0.7946   0.0304
  -6.750  -0.2926   0.08757   0.08407  -0.0204   0.7865   0.0309
  -6.500  -0.2828   0.08392   0.08037  -0.0201   0.7787   0.0315
  -6.250  -0.2699   0.08064   0.07704  -0.0216   0.7702   0.0322
  -6.000  -0.2556   0.07753   0.07384  -0.0237   0.7631   0.0332
  -5.750  -0.2383   0.07434   0.07060  -0.0265   0.7549   0.0344
  -5.500  -0.2191   0.07127   0.06741  -0.0294   0.7482   0.0362
  -5.250  -0.1716   0.07067   0.06649  -0.0396   0.7399   0.0390
  -5.000  -0.1522   0.06617   0.06184  -0.0419   0.7338   0.0394
  -4.750  -0.1437   0.06133   0.05704  -0.0411   0.7269   0.0402
  -4.500  -0.1284   0.05816   0.05381  -0.0410   0.7201   0.0412
  -4.250  -0.1079   0.05542   0.05099  -0.0421   0.7134   0.0427
  -4.000  -0.0839   0.05283   0.04827  -0.0436   0.7065   0.0451
  -3.750  -0.0343   0.05302   0.04795  -0.0481   0.7003   0.0502
  -3.500  -0.0157   0.04804   0.04291  -0.0492   0.6936   0.0511
  -3.250   0.0004   0.04485   0.03967  -0.0490   0.6881   0.0523
  -3.000   0.0223   0.04253   0.03731  -0.0494   0.6809   0.0543
  -2.750   0.0477   0.04052   0.03513  -0.0500   0.6748   0.0573
  -2.500   0.0896   0.03994   0.03401  -0.0514   0.6689   0.0638
  -2.250   0.1079   0.03642   0.03054  -0.0517   0.6624   0.0654
  -2.000   0.1305   0.03451   0.02851  -0.0517   0.6573   0.0677
  -1.750   0.1579   0.03295   0.02684  -0.0522   0.6500   0.0722
  -1.500   0.1913   0.03162   0.02505  -0.0524   0.6446   0.0787
  -1.250   0.2145   0.02963   0.02305  -0.0526   0.6389   0.0812
  -1.000   0.2418   0.02846   0.02176  -0.0528   0.6325   0.0877
  -0.750   0.2698   0.02703   0.02005  -0.0528   0.6277   0.0951
  -0.500   0.2972   0.02595   0.01889  -0.0530   0.6209   0.1016
  -0.250   0.3255   0.02471   0.01742  -0.0530   0.6154   0.1098
   0.000   0.3542   0.02403   0.01648  -0.0529   0.6105   0.1223
   0.250   0.3808   0.02279   0.01526  -0.0531   0.6041   0.1289
   0.500   0.4083   0.02189   0.01419  -0.0531   0.5992   0.1426
   0.750   0.4358   0.02121   0.01340  -0.0532   0.5934   0.1669
   1.000   0.4627   0.02037   0.01254  -0.0532   0.5877   0.1875
   1.250   0.5026   0.01791   0.00912  -0.0515   0.5840   0.0729
   1.500   0.5313   0.01714   0.00822  -0.0513   0.5787   0.0695
   1.750   0.5595   0.01655   0.00753  -0.0512   0.5731   0.0683
   2.000   0.5870   0.01613   0.00701  -0.0509   0.5688   0.0690
   2.250   0.6143   0.01585   0.00678  -0.0508   0.5633   0.0710
   2.500   0.6415   0.01564   0.00658  -0.0507   0.5582   0.0748
   2.750   0.6683   0.01542   0.00636  -0.0504   0.5541   0.0823
   3.000   0.6953   0.01535   0.00633  -0.0504   0.5488   0.0914
   3.250   0.7187   0.01441   0.00626  -0.0498   0.5439   0.4894
   3.500   0.7849   0.01366   0.00627  -0.0578   0.5390   1.0000
   3.750   0.8109   0.01389   0.00650  -0.0576   0.5338   1.0000
   4.000   0.8369   0.01407   0.00666  -0.0574   0.5288   1.0000
   4.250   0.8629   0.01426   0.00676  -0.0571   0.5249   1.0000
   4.500   0.8887   0.01453   0.00705  -0.0569   0.5206   1.0000
   4.750   0.9145   0.01477   0.00733  -0.0567   0.5155   1.0000
   5.000   0.9405   0.01496   0.00749  -0.0564   0.5114   1.0000
   5.250   0.9663   0.01521   0.00774  -0.0562   0.5072   1.0000
   5.500   0.9919   0.01548   0.00811  -0.0561   0.5016   1.0000
   5.750   1.0178   0.01565   0.00828  -0.0558   0.4969   1.0000
   6.000   1.0434   0.01583   0.00849  -0.0555   0.4910   1.0000
   6.250   1.0689   0.01568   0.00833  -0.0550   0.4810   1.0000
   6.500   1.0937   0.01555   0.00824  -0.0544   0.4688   1.0000
   6.750   1.1185   0.01541   0.00812  -0.0539   0.4558   1.0000
   7.000   1.1433   0.01539   0.00814  -0.0534   0.4439   1.0000
   7.250   1.1676   0.01531   0.00814  -0.0529   0.4278   1.0000
   7.500   1.1909   0.01524   0.00807  -0.0522   0.4021   1.0000
   7.750   1.2126   0.01542   0.00820  -0.0513   0.3608   1.0000
   8.000   1.2186   0.01743   0.00932  -0.0493   0.2148   1.0000
   8.250   1.2053   0.02139   0.01225  -0.0458   0.0595   1.0000
   8.500   1.2128   0.02301   0.01381  -0.0437   0.0423   1.0000
   8.750   1.2234   0.02424   0.01518  -0.0420   0.0388   1.0000
   9.000   1.2285   0.02571   0.01679  -0.0398   0.0365   1.0000
   9.250   1.2269   0.02762   0.01883  -0.0374   0.0351   1.0000
   9.500   1.2224   0.03024   0.02162  -0.0359   0.0337   1.0000
   9.750   1.2220   0.03275   0.02426  -0.0351   0.0328   1.0000
  10.000   1.2223   0.03532   0.02695  -0.0344   0.0318   1.0000
  10.250   1.2208   0.03812   0.02986  -0.0337   0.0311   1.0000
  10.500   1.2189   0.04096   0.03280  -0.0330   0.0305   1.0000
  10.750   1.2178   0.04371   0.03564  -0.0322   0.0300   1.0000
  11.000   1.2180   0.04625   0.03825  -0.0312   0.0295   1.0000
  11.250   1.2203   0.04854   0.04059  -0.0300   0.0291   1.0000
  11.500   1.2252   0.05046   0.04255  -0.0285   0.0287   1.0000
  11.750   1.2341   0.05194   0.04405  -0.0266   0.0283   1.0000
  12.000   1.2451   0.05328   0.04539  -0.0248   0.0274   1.0000
  12.250   1.2647   0.05388   0.04591  -0.0219   0.0260   1.0000
  12.500   1.2959   0.05425   0.04633  -0.0189   0.0258   1.0000
  12.750   1.3194   0.05551   0.04774  -0.0172   0.0261   1.0000
  13.000   1.3402   0.05721   0.04973  -0.0155   0.0273   1.0000
  13.250   1.1678   0.05667   0.04966  -0.0058   0.0264   1.0000
  13.500   1.1910   0.05835   0.05157  -0.0039   0.0272   1.0000
<< Back to NACA M20 AIRFOIL (m20-il)

Polar data table (+)

Polar graphs


<< Back to NACA M20 AIRFOIL (m20-il)