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

MH 108 11.98% (mh108-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: MH 108 11.98% (mh108-il)
Reynolds number: 50,000
Max Cl/Cd: 16.39 at α=3.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-mh108-il-50000.txt
Download as CSV file: xf-mh108-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: MH 108  11.98%                                  
  
 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
  ------ -------- --------- --------- -------- -------- --------
  -9.750  -0.6753   0.08691   0.08064  -0.0076   1.0000   0.1284
  -9.500  -0.6866   0.08080   0.07453  -0.0101   1.0000   0.1265
  -9.250  -0.7021   0.07524   0.06898  -0.0114   1.0000   0.1250
  -9.000  -0.7203   0.06968   0.06336  -0.0122   1.0000   0.1229
  -8.750  -0.7409   0.06379   0.05728  -0.0127   1.0000   0.1211
  -8.500  -0.7547   0.05841   0.05160  -0.0123   1.0000   0.1211
  -8.250  -0.7621   0.05354   0.04632  -0.0111   1.0000   0.1235
  -8.000  -0.7672   0.04868   0.04083  -0.0093   1.0000   0.1265
  -7.750  -0.7524   0.04584   0.03798  -0.0079   1.0000   0.1364
  -7.500  -0.7429   0.04249   0.03430  -0.0062   1.0000   0.1464
  -7.250  -0.7297   0.03982   0.03137  -0.0044   1.0000   0.1602
  -7.000  -0.7134   0.03740   0.02878  -0.0027   1.0000   0.1759
  -6.750  -0.6962   0.03534   0.02656  -0.0010   1.0000   0.1945
  -6.500  -0.6799   0.03322   0.02408   0.0008   1.0000   0.2139
  -6.250  -0.6580   0.03159   0.02259   0.0021   1.0000   0.2352
  -6.000  -0.6389   0.02997   0.02081   0.0037   1.0000   0.2590
  -5.750  -0.6174   0.02857   0.01945   0.0051   1.0000   0.2836
  -5.500  -0.5962   0.02726   0.01821   0.0066   1.0000   0.3105
  -5.250  -0.5754   0.02607   0.01708   0.0082   1.0000   0.3402
  -5.000  -0.5547   0.02494   0.01602   0.0098   1.0000   0.3722
  -4.750  -0.5344   0.02393   0.01510   0.0115   1.0000   0.4073
  -4.500  -0.5142   0.02302   0.01438   0.0134   1.0000   0.4449
  -4.250  -0.4947   0.02215   0.01371   0.0155   1.0000   0.4854
  -4.000  -0.4761   0.02139   0.01317   0.0178   1.0000   0.5297
  -3.750  -0.4583   0.02073   0.01276   0.0203   1.0000   0.5770
  -3.500  -0.4424   0.02019   0.01242   0.0232   1.0000   0.6282
  -3.250  -0.4258   0.01986   0.01238   0.0263   1.0000   0.6803
  -3.000  -0.4125   0.01974   0.01247   0.0299   1.0000   0.7352
  -2.750  -0.3931   0.01996   0.01285   0.0327   1.0000   0.7943
  -2.500  -0.3380   0.02093   0.01375   0.0302   1.0000   0.8612
  -2.250  -0.2021   0.02223   0.01459   0.0137   1.0000   0.9239
  -2.000  -0.0682   0.02211   0.01413  -0.0065   1.0000   0.9743
  -1.750   0.0351   0.02138   0.01320  -0.0251   0.9739   1.0000
  -1.500   0.1257   0.02061   0.01214  -0.0391   0.9179   1.0000
  -1.250   0.1651   0.02041   0.01167  -0.0422   0.8693   1.0000
  -1.000   0.1816   0.02053   0.01156  -0.0408   0.8323   1.0000
  -0.750   0.1953   0.02073   0.01158  -0.0388   0.8016   1.0000
  -0.500   0.2095   0.02095   0.01163  -0.0366   0.7761   1.0000
  -0.250   0.2255   0.02123   0.01179  -0.0351   0.7519   1.0000
   0.000   0.2421   0.02151   0.01192  -0.0332   0.7319   1.0000
   0.250   0.2597   0.02190   0.01222  -0.0320   0.7115   1.0000
   0.500   0.2779   0.02231   0.01254  -0.0306   0.6938   1.0000
   0.750   0.2970   0.02276   0.01290  -0.0293   0.6775   1.0000
   1.000   0.3160   0.02327   0.01333  -0.0281   0.6623   1.0000
   1.250   0.3354   0.02384   0.01384  -0.0270   0.6481   1.0000
   1.500   0.3550   0.02446   0.01442  -0.0259   0.6346   1.0000
   1.750   0.3748   0.02513   0.01506  -0.0249   0.6219   1.0000
   2.000   0.3946   0.02580   0.01569  -0.0237   0.6105   1.0000
   2.250   0.4146   0.02646   0.01631  -0.0225   0.5996   1.0000
   2.500   0.4342   0.02740   0.01729  -0.0219   0.5876   1.0000
   2.750   0.4537   0.02838   0.01829  -0.0212   0.5768   1.0000
   3.000   0.4738   0.02911   0.01900  -0.0198   0.5676   1.0000
   3.250   0.4927   0.03019   0.02015  -0.0192   0.5567   1.0000
   3.500   0.5106   0.03149   0.02153  -0.0186   0.5463   1.0000
   3.750   0.5305   0.03236   0.02240  -0.0173   0.5375   1.0000
   4.000   0.5479   0.03365   0.02377  -0.0166   0.5271   1.0000
   4.250   0.5631   0.03534   0.02556  -0.0161   0.5170   1.0000
   4.500   0.5834   0.03614   0.02636  -0.0145   0.5082   1.0000
   4.750   0.5970   0.03799   0.02835  -0.0140   0.4978   1.0000
   5.000   0.6085   0.04005   0.03052  -0.0134   0.4876   1.0000
   5.250   0.6299   0.04076   0.03123  -0.0116   0.4787   1.0000
   5.500   0.6375   0.04325   0.03386  -0.0112   0.4681   1.0000
   5.750   0.6416   0.04609   0.03680  -0.0108   0.4579   1.0000
   6.000   0.6625   0.04699   0.03774  -0.0090   0.4486   1.0000
   6.250   0.6638   0.05011   0.04098  -0.0087   0.4384   1.0000
   6.500   0.6537   0.05445   0.04539  -0.0091   0.4293   1.0000
   6.750   0.6779   0.05511   0.04612  -0.0071   0.4192   1.0000
   7.000   0.6376   0.06245   0.05342  -0.0095   0.4135   1.0000
   7.250   0.6532   0.06421   0.05525  -0.0082   0.4036   1.0000
   7.500   0.6187   0.07020   0.06113  -0.0093   0.4022   1.0000
   7.750   0.5995   0.07499   0.06587  -0.0101   0.4009   1.0000
   8.000   0.5880   0.07935   0.07021  -0.0110   0.3999   1.0000
   8.250   0.5822   0.08378   0.07466  -0.0122   0.4024   1.0000
   8.500   0.6080   0.08598   0.07697  -0.0113   0.3910   1.0000
   8.750   0.5313   0.09890   0.08973  -0.0225   0.4944   1.0000
   9.000   0.5411   0.10173   0.09263  -0.0221   0.4796   1.0000
<< Back to MH 108 11.98% (mh108-il)

Polar data table (+)

Polar graphs


<< Back to MH 108 11.98% (mh108-il)