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

MH 82 13.31% (mh82-il) Xfoil prediction polar at RE=50,000 Ncrit=5


Details Polar file
Airfoil: MH 82 13.31% (mh82-il)
Reynolds number: 50,000
Max Cl/Cd: 23.42 at α=5°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-mh82-il-50000-n5.txt
Download as CSV file: xf-mh82-il-50000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: MH 82  13.31%                                   
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.050 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.500  -0.3330   0.09712   0.09163  -0.0089   1.0000   0.0546
  -8.250  -0.3283   0.09281   0.08745  -0.0120   1.0000   0.0524
  -8.000  -0.3350   0.08642   0.08117  -0.0192   1.0000   0.0487
  -7.750  -0.3256   0.08242   0.07726  -0.0231   0.9443   0.0479
  -7.500  -0.3166   0.07752   0.07229  -0.0287   0.8913   0.0469
  -7.250  -0.3139   0.07308   0.06774  -0.0322   0.8590   0.0459
  -7.000  -0.3147   0.06839   0.06289  -0.0350   0.8353   0.0449
  -6.750  -0.3162   0.06328   0.05754  -0.0376   0.8161   0.0438
  -6.250  -0.3126   0.05291   0.04626  -0.0408   0.7838   0.0418
  -6.000  -0.3018   0.04940   0.04236  -0.0407   0.7679   0.0417
  -5.750  -0.2883   0.04608   0.03859  -0.0404   0.7528   0.0416
  -5.500  -0.2728   0.04283   0.03481  -0.0399   0.7384   0.0420
  -5.250  -0.2538   0.04112   0.03294  -0.0391   0.7235   0.0435
  -5.000  -0.2341   0.03905   0.03050  -0.0383   0.7100   0.0453
  -4.750  -0.2130   0.03678   0.02770  -0.0373   0.6973   0.0470
  -4.500  -0.1893   0.03457   0.02493  -0.0365   0.6836   0.0485
  -4.250  -0.1649   0.03277   0.02279  -0.0357   0.6706   0.0500
  -4.000  -0.1405   0.03144   0.02126  -0.0347   0.6584   0.0526
  -3.750  -0.1146   0.03015   0.01957  -0.0337   0.6462   0.0571
  -3.500  -0.0890   0.02907   0.01846  -0.0331   0.6334   0.0619
  -3.250  -0.0627   0.02800   0.01716  -0.0321   0.6220   0.0684
  -3.000  -0.0365   0.02712   0.01612  -0.0312   0.6109   0.0788
  -2.750  -0.0098   0.02627   0.01529  -0.0307   0.5988   0.0955
  -2.500   0.0163   0.02535   0.01439  -0.0301   0.5887   0.1300
  -2.250   0.0406   0.02436   0.01371  -0.0294   0.5780   0.2010
  -2.000   0.0625   0.02334   0.01331  -0.0286   0.5677   0.3172
  -1.750   0.0790   0.02212   0.01308  -0.0261   0.5589   0.5200
  -1.500   0.1213   0.02137   0.01320  -0.0251   0.5471   0.8578
  -1.250   0.1951   0.02138   0.01263  -0.0321   0.5339   0.9801
  -1.000   0.2348   0.02142   0.01222  -0.0344   0.5240   1.0000
  -0.750   0.2576   0.02168   0.01221  -0.0338   0.5144   1.0000
  -0.500   0.2800   0.02186   0.01205  -0.0328   0.5068   1.0000
  -0.250   0.3031   0.02219   0.01216  -0.0322   0.4976   1.0000
   0.000   0.3260   0.02242   0.01210  -0.0313   0.4903   1.0000
   0.250   0.3492   0.02281   0.01229  -0.0307   0.4819   1.0000
   0.500   0.3724   0.02311   0.01236  -0.0298   0.4747   1.0000
   0.750   0.3957   0.02349   0.01254  -0.0291   0.4675   1.0000
   1.000   0.4191   0.02390   0.01280  -0.0285   0.4598   1.0000
   1.250   0.4428   0.02416   0.01279  -0.0275   0.4546   1.0000
   1.500   0.4661   0.02481   0.01340  -0.0273   0.4463   1.0000
   1.750   0.4899   0.02516   0.01356  -0.0265   0.4404   1.0000
   2.000   0.5135   0.02569   0.01398  -0.0260   0.4341   1.0000
   2.250   0.5368   0.02629   0.01451  -0.0256   0.4273   1.0000
   2.500   0.5612   0.02665   0.01468  -0.0248   0.4226   1.0000
   2.750   0.5840   0.02742   0.01544  -0.0246   0.4161   1.0000
   3.000   0.6072   0.02806   0.01602  -0.0241   0.4101   1.0000
   3.250   0.6320   0.02841   0.01621  -0.0234   0.4058   1.0000
   3.500   0.6535   0.02940   0.01724  -0.0232   0.3995   1.0000
   3.750   0.6760   0.03019   0.01802  -0.0229   0.3940   1.0000
   4.000   0.7006   0.03062   0.01833  -0.0223   0.3900   1.0000
   4.250   0.7216   0.03163   0.01937  -0.0220   0.3846   1.0000
   4.500   0.7416   0.03276   0.02056  -0.0218   0.3788   1.0000
   4.750   0.7653   0.03334   0.02109  -0.0213   0.3748   1.0000
   5.000   0.7906   0.03376   0.02138  -0.0206   0.3717   1.0000
   5.250   0.8023   0.03593   0.02380  -0.0208   0.3648   1.0000
   5.500   0.8227   0.03689   0.02477  -0.0203   0.3603   1.0000
   5.750   0.8475   0.03733   0.02513  -0.0197   0.3571   1.0000
   6.000   0.8580   0.03950   0.02748  -0.0197   0.3517   1.0000
   6.250   0.8673   0.04171   0.02982  -0.0195   0.3463   1.0000
   6.500   0.8882   0.04254   0.03065  -0.0190   0.3428   1.0000
   6.750   0.9154   0.04273   0.03075  -0.0183   0.3402   1.0000
   7.000   0.8819   0.04933   0.03771  -0.0191   0.3312   1.0000
   7.250   0.8941   0.05102   0.03946  -0.0188   0.3275   1.0000
   7.500   0.9198   0.05129   0.03970  -0.0180   0.3251   1.0000
   8.000   0.8460   0.06582   0.05443  -0.0220   0.3095   1.0000
   8.500   0.8006   0.07913   0.06779  -0.0270   0.2956   1.0000
   8.750   0.8230   0.07967   0.06836  -0.0261   0.2935   1.0000
   9.250   0.7878   0.09177   0.08054  -0.0307   0.2816   1.0000
<< Back to MH 82 13.31% (mh82-il)

Polar data table (+)

Polar graphs


<< Back to MH 82 13.31% (mh82-il)