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MH 20 9.01% (mh20-il) Xfoil prediction polar at RE=100,000 Ncrit=9


Details Polar file
Airfoil: MH 20 9.01% (mh20-il)
Reynolds number: 100,000
Max Cl/Cd: 50.66 at α=6.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-mh20-il-100000.txt
Download as CSV file: xf-mh20-il-100000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: MH 20  9.01%                                    
  
 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
  ------ -------- --------- --------- -------- -------- --------
  -9.250  -0.4164   0.10286   0.09874  -0.0010   1.0000   0.0838
  -9.000  -0.5039   0.10432   0.09997   0.0002   1.0000   0.0787
  -8.750  -0.4967   0.10074   0.09634  -0.0006   1.0000   0.0815
  -8.500  -0.4346   0.08992   0.08599  -0.0102   1.0000   0.0899
  -8.250  -0.4170   0.08574   0.08180  -0.0084   1.0000   0.0930
  -8.000  -0.4148   0.08166   0.07776  -0.0099   1.0000   0.0964
  -7.750  -0.4256   0.07740   0.07359  -0.0142   1.0000   0.1003
  -7.500  -0.4545   0.07298   0.06923  -0.0228   1.0000   0.1017
  -7.250  -0.4452   0.06747   0.06379  -0.0223   1.0000   0.1037
  -7.000  -0.4316   0.06359   0.05994  -0.0212   1.0000   0.1070
  -6.750  -0.4904   0.06840   0.06435  -0.0237   1.0000   0.1057
  -6.500  -0.4843   0.06473   0.06057  -0.0272   1.0000   0.1141
  -6.250  -0.4759   0.06022   0.05604  -0.0284   1.0000   0.1180
  -5.000  -0.4645   0.03660   0.03065  -0.0207   0.9964   0.0440
  -4.750  -0.4286   0.03232   0.02520  -0.0215   0.9883   0.0372
  -4.500  -0.3936   0.02839   0.02074  -0.0233   0.9820   0.0362
  -4.250  -0.3591   0.02568   0.01749  -0.0245   0.9748   0.0365
  -4.000  -0.3210   0.02291   0.01448  -0.0267   0.9694   0.0412
  -3.750  -0.2843   0.02147   0.01274  -0.0280   0.9625   0.0494
  -3.500  -0.2471   0.01921   0.01061  -0.0296   0.9571   0.0687
  -3.250  -0.2175   0.01737   0.00941  -0.0302   0.9499   0.1917
  -3.000  -0.1864   0.01595   0.00872  -0.0313   0.9434   0.3301
  -2.750  -0.0863   0.01376   0.00848  -0.0418   0.9516   0.9210
  -2.500   0.0255   0.01368   0.00778  -0.0571   0.9587   1.0000
  -2.250   0.0646   0.01363   0.00742  -0.0601   0.9486   1.0000
  -2.000   0.1004   0.01362   0.00713  -0.0623   0.9386   1.0000
  -1.750   0.1331   0.01366   0.00695  -0.0636   0.9283   1.0000
  -1.500   0.1584   0.01378   0.00691  -0.0636   0.9167   1.0000
  -1.250   0.1815   0.01394   0.00693  -0.0630   0.9055   1.0000
  -1.000   0.2039   0.01412   0.00695  -0.0621   0.8947   1.0000
  -0.750   0.2255   0.01431   0.00702  -0.0610   0.8844   1.0000
  -0.500   0.2466   0.01449   0.00710  -0.0596   0.8748   1.0000
  -0.250   0.2669   0.01473   0.00726  -0.0584   0.8639   1.0000
   0.000   0.2875   0.01498   0.00744  -0.0571   0.8534   1.0000
   0.250   0.3081   0.01521   0.00761  -0.0556   0.8436   1.0000
   0.500   0.3283   0.01539   0.00772  -0.0538   0.8346   1.0000
   0.750   0.3492   0.01565   0.00794  -0.0525   0.8235   1.0000
   1.000   0.3701   0.01590   0.00816  -0.0511   0.8129   1.0000
   1.250   0.3910   0.01613   0.00836  -0.0496   0.8025   1.0000
   1.500   0.4117   0.01629   0.00851  -0.0478   0.7929   1.0000
   1.750   0.4327   0.01647   0.00868  -0.0462   0.7821   1.0000
   2.000   0.4541   0.01669   0.00890  -0.0447   0.7704   1.0000
   2.250   0.4754   0.01688   0.00910  -0.0432   0.7587   1.0000
   2.500   0.4969   0.01704   0.00927  -0.0416   0.7469   1.0000
   2.750   0.5185   0.01717   0.00942  -0.0399   0.7351   1.0000
   3.000   0.5402   0.01726   0.00960  -0.0382   0.7231   1.0000
   3.250   0.5622   0.01732   0.00970  -0.0365   0.7108   1.0000
   3.500   0.5845   0.01737   0.00979  -0.0348   0.6977   1.0000
   3.750   0.6069   0.01741   0.00988  -0.0332   0.6838   1.0000
   4.000   0.6294   0.01743   0.01004  -0.0316   0.6689   1.0000
   4.250   0.6522   0.01741   0.01009  -0.0300   0.6533   1.0000
   4.500   0.6753   0.01733   0.01007  -0.0283   0.6373   1.0000
   4.750   0.6987   0.01718   0.00998  -0.0266   0.6210   1.0000
   5.000   0.7218   0.01712   0.01003  -0.0251   0.6012   1.0000
   5.250   0.7454   0.01694   0.00991  -0.0234   0.5813   1.0000
   5.500   0.7687   0.01683   0.00992  -0.0219   0.5578   1.0000
   5.750   0.7919   0.01670   0.00997  -0.0204   0.5319   1.0000
   6.000   0.8147   0.01660   0.00992  -0.0187   0.5020   1.0000
   6.250   0.8363   0.01665   0.00998  -0.0170   0.4645   1.0000
   6.500   0.8561   0.01690   0.01012  -0.0152   0.4187   1.0000
   6.750   0.8738   0.01742   0.01056  -0.0135   0.3595   1.0000
   7.000   0.8879   0.01835   0.01121  -0.0116   0.2765   1.0000
   7.250   0.8868   0.02134   0.01302  -0.0085   0.1233   1.0000
   7.500   0.8883   0.02433   0.01545  -0.0053   0.0683   1.0000
   7.750   0.8948   0.02675   0.01771  -0.0024   0.0512   1.0000
   8.000   0.9072   0.02870   0.01966  -0.0003   0.0407   1.0000
   8.250   0.9261   0.03178   0.02272   0.0015   0.0364   1.0000
   8.500   0.9478   0.03414   0.02533   0.0029   0.0335   1.0000
   8.750   0.9636   0.03625   0.02758   0.0043   0.0294   1.0000
   9.000   0.9795   0.04064   0.03212   0.0051   0.0270   1.0000
   9.250   0.9913   0.04474   0.03660   0.0067   0.0268   1.0000
   9.500   1.0002   0.04800   0.04036   0.0085   0.0269   1.0000
   9.750   1.0076   0.05039   0.04309   0.0104   0.0273   1.0000
  10.000   1.0089   0.05315   0.04625   0.0127   0.0279   1.0000
  10.250   1.0020   0.05639   0.04997   0.0153   0.0290   1.0000
  10.500   0.9821   0.05971   0.05366   0.0182   0.0297   1.0000
  10.750   0.9609   0.06355   0.05780   0.0196   0.0305   1.0000
  11.000   0.9410   0.06778   0.06227   0.0194   0.0314   1.0000
  11.250   0.9201   0.07261   0.06729   0.0181   0.0318   1.0000
  11.500   0.8981   0.07816   0.07301   0.0156   0.0320   1.0000
  11.750   0.8770   0.08429   0.07928   0.0120   0.0323   1.0000
  12.000   0.8538   0.09175   0.08673   0.0071   0.0326   1.0000
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