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

MH 81 13% (mh81-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: MH 81 13% (mh81-il)
Reynolds number: 50,000
Max Cl/Cd: 11.83 at α=1°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-mh81-il-50000.txt
Download as CSV file: xf-mh81-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: MH 81  13%                                      
  
 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
  ------ -------- --------- --------- -------- -------- --------
  -8.500  -0.3863   0.10981   0.10452   0.0118   1.0000   0.2346
  -8.250  -0.3588   0.10412   0.09888   0.0126   1.0000   0.2419
  -8.000  -0.3844   0.10394   0.09891   0.0083   1.0000   0.2527
  -7.750  -0.3431   0.09751   0.09251   0.0103   1.0000   0.2623
  -7.500  -0.3453   0.09466   0.08982   0.0081   1.0000   0.2736
  -7.250  -0.3633   0.09367   0.08905   0.0046   1.0000   0.2879
  -7.000  -0.3338   0.08894   0.08441   0.0050   1.0000   0.3042
  -6.750  -0.3274   0.08604   0.08169   0.0033   1.0000   0.3223
  -6.500  -0.3240   0.08356   0.07941   0.0015   1.0000   0.3405
  -6.250  -0.3173   0.08141   0.07744   0.0025   1.0000   0.3571
  -6.000  -0.2850   0.07723   0.07330  -0.0004   0.9795   0.3818
  -5.750  -0.3195   0.05891   0.05358  -0.0447   0.9449   0.1505
  -5.500  -0.2816   0.05205   0.04647  -0.0497   0.9241   0.1311
  -5.250  -0.2591   0.04698   0.04050  -0.0522   0.9027   0.1185
  -5.000  -0.2332   0.04364   0.03684  -0.0526   0.8820   0.1166
  -4.750  -0.2128   0.04095   0.03372  -0.0517   0.8616   0.1146
  -4.500  -0.1929   0.03864   0.03094  -0.0502   0.8417   0.1133
  -4.250  -0.1718   0.03672   0.02853  -0.0485   0.8231   0.1152
  -4.000  -0.1496   0.03525   0.02644  -0.0466   0.8053   0.1195
  -3.750  -0.1269   0.03364   0.02484  -0.0452   0.7880   0.1270
  -3.500  -0.1022   0.03231   0.02321  -0.0436   0.7718   0.1376
  -3.250  -0.0774   0.03114   0.02198  -0.0420   0.7560   0.1559
  -3.000  -0.0519   0.02989   0.02080  -0.0403   0.7412   0.1902
  -2.750  -0.0267   0.02806   0.01943  -0.0384   0.7279   0.2850
  -2.500  -0.0167   0.02584   0.01888  -0.0347   0.7144   0.5049
  -2.250   0.1575   0.02497   0.01812  -0.0492   0.6873   1.0000
  -2.000   0.1722   0.02515   0.01792  -0.0476   0.6754   1.0000
  -1.750   0.1914   0.02559   0.01809  -0.0472   0.6613   1.0000
  -1.500   0.2106   0.02613   0.01836  -0.0467   0.6486   1.0000
  -1.250   0.2278   0.02647   0.01838  -0.0447   0.6387   1.0000
  -1.000   0.2482   0.02716   0.01888  -0.0444   0.6259   1.0000
  -0.750   0.2681   0.02793   0.01945  -0.0440   0.6148   1.0000
  -0.500   0.2868   0.02840   0.01966  -0.0422   0.6058   1.0000
  -0.250   0.3070   0.02947   0.02061  -0.0424   0.5942   1.0000
   0.000   0.3258   0.02985   0.02072  -0.0401   0.5868   1.0000
   0.250   0.3453   0.03131   0.02214  -0.0410   0.5755   1.0000
   0.500   0.3649   0.03172   0.02230  -0.0387   0.5690   1.0000
   0.750   0.3819   0.03356   0.02414  -0.0398   0.5581   1.0000
   1.000   0.4022   0.03399   0.02434  -0.0376   0.5519   1.0000
   1.250   0.4152   0.03640   0.02678  -0.0390   0.5423   1.0000
   1.500   0.4349   0.03718   0.02741  -0.0375   0.5360   1.0000
   1.750   0.4451   0.03959   0.02981  -0.0380   0.5283   1.0000
   2.000   0.4580   0.04138   0.03154  -0.0376   0.5212   1.0000
   2.250   0.4828   0.04164   0.03161  -0.0353   0.5165   1.0000
   2.500   0.4709   0.04662   0.03670  -0.0379   0.5105   1.0000
   2.750   0.4727   0.04952   0.03957  -0.0381   0.5055   1.0000
   3.000   0.5037   0.04961   0.03951  -0.0360   0.4999   1.0000
   3.250   0.4952   0.05383   0.04373  -0.0373   0.4975   1.0000
   3.500   0.4862   0.05787   0.04775  -0.0382   0.4974   1.0000
   3.750   0.4813   0.06134   0.05117  -0.0386   0.4978   1.0000
   4.000   0.4804   0.06449   0.05428  -0.0386   0.4995   1.0000
   4.750   0.3592   0.07931   0.06921  -0.0448   0.6266   1.0000
   5.000   0.3718   0.08165   0.07148  -0.0446   0.6179   1.0000
   5.250   0.3902   0.08391   0.07369  -0.0447   0.6071   1.0000
   5.500   0.3938   0.08569   0.07541  -0.0439   0.5975   1.0000
   5.750   0.4233   0.08892   0.07859  -0.0448   0.5880   1.0000
   6.000   0.4194   0.09002   0.07964  -0.0435   0.5762   1.0000
   6.250   0.4562   0.09426   0.08385  -0.0450   0.5693   1.0000
   6.500   0.4436   0.09462   0.08417  -0.0433   0.5580   1.0000
   6.750   0.4797   0.09902   0.08855  -0.0447   0.5518   1.0000
   7.000   0.4676   0.09935   0.08885  -0.0432   0.5397   1.0000
   7.500   0.4907   0.10418   0.09364  -0.0431   0.5219   1.0000
   7.750   0.5251   0.10887   0.09834  -0.0444   0.5168   1.0000
   8.000   0.5093   0.10905   0.09849  -0.0432   0.5062   1.0000
   8.250   0.5434   0.11359   0.10305  -0.0442   0.4997   1.0000
<< Back to MH 81 13% (mh81-il)

Polar data table (+)

Polar graphs


<< Back to MH 81 13% (mh81-il)