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MH 62 9.3% (mh62-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: MH 62 9.3% (mh62-il)
Reynolds number: 50,000
Max Cl/Cd: 28.73 at α=6.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-mh62-il-50000.txt
Download as CSV file: xf-mh62-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: MH 62  9.3%                                     
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.050 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.000  -0.4248   0.10663   0.10085   0.0118   1.0000   0.2433
  -8.750  -0.4260   0.10333   0.09760   0.0109   1.0000   0.2530
  -8.500  -0.5262   0.10726   0.10114   0.0154   1.0000   0.2539
  -8.250  -0.5045   0.10178   0.09567   0.0161   1.0000   0.2606
  -7.750  -0.5082   0.09574   0.08980   0.0136   1.0000   0.2810
  -7.500  -0.4966   0.09105   0.08515   0.0132   1.0000   0.2863
  -7.250  -0.5055   0.08844   0.08267   0.0116   1.0000   0.2972
  -7.000  -0.4911   0.08380   0.07806   0.0114   1.0000   0.3014
  -6.750  -0.5314   0.06626   0.06041  -0.0196   1.0000   0.1365
  -6.500  -0.5305   0.05749   0.05119  -0.0261   1.0000   0.1133
  -6.250  -0.5231   0.05065   0.04369  -0.0289   1.0000   0.1017
  -6.000  -0.5073   0.04597   0.03876  -0.0292   1.0000   0.0977
  -5.750  -0.4919   0.04094   0.03305  -0.0294   1.0000   0.0929
  -5.500  -0.4731   0.03702   0.02819  -0.0287   1.0000   0.0905
  -5.250  -0.4522   0.03379   0.02478  -0.0279   1.0000   0.0929
  -5.000  -0.4315   0.03144   0.02212  -0.0268   1.0000   0.1002
  -4.750  -0.4109   0.02893   0.01936  -0.0255   1.0000   0.1078
  -4.500  -0.3901   0.02670   0.01688  -0.0237   1.0000   0.1188
  -4.250  -0.3719   0.02498   0.01514  -0.0218   1.0000   0.1430
  -4.000  -0.3549   0.02312   0.01360  -0.0197   1.0000   0.1875
  -3.750  -0.3407   0.02096   0.01223  -0.0169   1.0000   0.3024
  -3.500  -0.3332   0.01987   0.01180  -0.0129   1.0000   0.4344
  -3.250  -0.3240   0.01925   0.01160  -0.0088   1.0000   0.5300
  -3.000  -0.3131   0.01862   0.01123  -0.0046   1.0000   0.6186
  -2.750  -0.3019   0.01797   0.01094   0.0000   1.0000   0.7091
  -2.500  -0.2810   0.01742   0.01075   0.0042   1.0000   0.8350
  -2.250  -0.1070   0.01699   0.00919  -0.0225   1.0000   1.0000
  -2.000  -0.1216   0.01687   0.00884  -0.0182   1.0000   1.0000
  -1.750  -0.1191   0.01694   0.00862  -0.0156   1.0000   1.0000
  -1.500  -0.1105   0.01711   0.00851  -0.0137   1.0000   1.0000
  -1.250  -0.0989   0.01733   0.00847  -0.0122   1.0000   1.0000
  -1.000  -0.0856   0.01760   0.00851  -0.0110   1.0000   1.0000
  -0.750  -0.0712   0.01792   0.00862  -0.0099   1.0000   1.0000
  -0.500  -0.0562   0.01829   0.00877  -0.0091   1.0000   1.0000
  -0.250  -0.0407   0.01871   0.00901  -0.0084   1.0000   1.0000
   0.000  -0.0251   0.01917   0.00933  -0.0078   1.0000   1.0000
   0.250   0.0332   0.02021   0.01019  -0.0152   0.9832   1.0000
   0.500   0.0890   0.02116   0.01101  -0.0219   0.9648   1.0000
   0.750   0.1447   0.02206   0.01183  -0.0283   0.9457   1.0000
   1.000   0.2010   0.02288   0.01262  -0.0344   0.9265   1.0000
   1.250   0.2513   0.02359   0.01332  -0.0392   0.9057   1.0000
   1.500   0.3102   0.02418   0.01397  -0.0451   0.8859   1.0000
   1.750   0.3502   0.02475   0.01459  -0.0473   0.8641   1.0000
   2.000   0.3918   0.02522   0.01512  -0.0493   0.8432   1.0000
   2.250   0.4253   0.02569   0.01564  -0.0497   0.8214   1.0000
   2.500   0.4543   0.02614   0.01615  -0.0492   0.7997   1.0000
   2.750   0.4840   0.02644   0.01655  -0.0483   0.7787   1.0000
   3.000   0.5071   0.02687   0.01704  -0.0466   0.7562   1.0000
   3.250   0.5343   0.02699   0.01722  -0.0447   0.7360   1.0000
   3.500   0.5553   0.02738   0.01767  -0.0426   0.7128   1.0000
   3.750   0.5813   0.02726   0.01765  -0.0400   0.6933   1.0000
   4.000   0.6019   0.02758   0.01806  -0.0378   0.6693   1.0000
   4.250   0.6245   0.02766   0.01821  -0.0354   0.6468   1.0000
   4.500   0.6486   0.02744   0.01805  -0.0325   0.6253   1.0000
   4.750   0.6707   0.02755   0.01829  -0.0302   0.6005   1.0000
   5.000   0.6936   0.02753   0.01833  -0.0276   0.5761   1.0000
   5.250   0.7174   0.02740   0.01823  -0.0250   0.5517   1.0000
   5.500   0.7415   0.02726   0.01809  -0.0224   0.5265   1.0000
   5.750   0.7649   0.02737   0.01819  -0.0201   0.4993   1.0000
   6.000   0.7876   0.02774   0.01866  -0.0181   0.4704   1.0000
   6.250   0.8101   0.02830   0.01925  -0.0163   0.4411   1.0000
   6.500   0.8323   0.02897   0.01994  -0.0145   0.4113   1.0000
   6.750   0.8545   0.02978   0.02077  -0.0129   0.3822   1.0000
   7.000   0.8764   0.03073   0.02173  -0.0113   0.3535   1.0000
   7.250   0.8978   0.03179   0.02281  -0.0098   0.3252   1.0000
   7.500   0.9187   0.03303   0.02417  -0.0084   0.2978   1.0000
   7.750   0.9388   0.03443   0.02562  -0.0070   0.2713   1.0000
   8.000   0.9581   0.03585   0.02707  -0.0055   0.2446   1.0000
   8.250   0.9772   0.03763   0.02887  -0.0041   0.2200   1.0000
   8.500   0.9952   0.03921   0.03039  -0.0027   0.1944   1.0000
   8.750   1.0083   0.04202   0.03353  -0.0013   0.1746   1.0000
   9.000   1.0245   0.04405   0.03548   0.0001   0.1536   1.0000
   9.250   1.0356   0.04749   0.03919   0.0014   0.1401   1.0000
   9.500   1.0438   0.05045   0.04236   0.0028   0.1267   1.0000
   9.750   1.0408   0.05473   0.04719   0.0040   0.1204   1.0000
  10.000   1.0485   0.05894   0.05149   0.0050   0.1142   1.0000
  10.250   1.0325   0.06405   0.05706   0.0057   0.1133   1.0000
  10.500   1.0135   0.06914   0.06244   0.0059   0.1132   1.0000
  10.750   0.9923   0.07411   0.06755   0.0059   0.1139   1.0000
  11.000   0.9710   0.07958   0.07310   0.0046   0.1147   1.0000
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