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

MH 22 7.2% (mh22-il) Xfoil prediction polar at RE=200,000 Ncrit=9


Details Polar file
Airfoil: MH 22 7.2% (mh22-il)
Reynolds number: 200,000
Max Cl/Cd: 63.9 at α=4.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-mh22-il-200000.txt
Download as CSV file: xf-mh22-il-200000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: MH 22  7.2%                                     
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.200 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.750  -0.5238   0.10382   0.10047   0.0014   1.0000   0.0273
  -8.500  -0.5210   0.09985   0.09654  -0.0010   1.0000   0.0273
  -8.250  -0.5262   0.09361   0.09038  -0.0032   1.0000   0.0279
  -8.000  -0.5230   0.08928   0.08607  -0.0025   1.0000   0.0288
  -7.750  -0.5196   0.08586   0.08269  -0.0033   1.0000   0.0295
  -7.500  -0.5184   0.08244   0.07931  -0.0048   1.0000   0.0300
  -7.250  -0.5166   0.07863   0.07553  -0.0073   1.0000   0.0306
  -7.000  -0.5118   0.07437   0.07129  -0.0108   1.0000   0.0313
  -6.750  -0.5050   0.06992   0.06682  -0.0143   1.0000   0.0322
  -6.500  -0.4962   0.06521   0.06208  -0.0179   1.0000   0.0332
  -6.250  -0.4851   0.06027   0.05706  -0.0214   1.0000   0.0346
  -6.000  -0.4711   0.05528   0.05192  -0.0245   1.0000   0.0363
  -5.750  -0.4480   0.05162   0.04769  -0.0275   1.0000   0.0395
  -5.500  -0.4414   0.04421   0.04019  -0.0287   1.0000   0.0410
  -5.250  -0.4252   0.04071   0.03665  -0.0286   1.0000   0.0425
  -5.000  -0.4069   0.03735   0.03302  -0.0285   1.0000   0.0443
  -4.750  -0.3845   0.02737   0.02195  -0.0258   1.0000   0.0177
  -4.500  -0.3601   0.02569   0.02001  -0.0246   1.0000   0.0162
  -4.250  -0.3401   0.02214   0.01587  -0.0232   1.0000   0.0158
  -4.000  -0.3197   0.01924   0.01248  -0.0217   1.0000   0.0178
  -3.750  -0.2998   0.01827   0.01149  -0.0208   1.0000   0.0232
  -3.500  -0.2816   0.01660   0.00959  -0.0187   1.0000   0.0233
  -3.250  -0.2510   0.01490   0.00770  -0.0192   0.9941   0.0238
  -3.000  -0.2079   0.01331   0.00598  -0.0219   0.9826   0.0261
  -2.750  -0.1681   0.01117   0.00413  -0.0241   0.9687   0.1058
  -2.500  -0.1315   0.01004   0.00386  -0.0267   0.9519   0.2938
  -2.250  -0.1018   0.00892   0.00358  -0.0273   0.9331   0.5195
  -2.000  -0.0571   0.00771   0.00357  -0.0291   0.9237   0.9043
  -1.750   0.0312   0.00767   0.00319  -0.0411   0.9227   1.0000
  -1.500   0.0585   0.00768   0.00297  -0.0408   0.8967   1.0000
  -1.250   0.0805   0.00774   0.00281  -0.0394   0.8756   1.0000
  -1.000   0.1027   0.00782   0.00271  -0.0380   0.8569   1.0000
  -0.750   0.1253   0.00791   0.00257  -0.0368   0.8409   1.0000
  -0.500   0.1484   0.00801   0.00251  -0.0357   0.8264   1.0000
  -0.250   0.1720   0.00811   0.00247  -0.0347   0.8129   1.0000
   0.000   0.1959   0.00821   0.00245  -0.0337   0.7999   1.0000
   0.250   0.2201   0.00831   0.00245  -0.0329   0.7874   1.0000
   0.500   0.2445   0.00842   0.00247  -0.0321   0.7751   1.0000
   0.750   0.2691   0.00853   0.00251  -0.0313   0.7631   1.0000
   1.000   0.2938   0.00864   0.00255  -0.0305   0.7510   1.0000
   1.250   0.3186   0.00876   0.00260  -0.0297   0.7390   1.0000
   1.500   0.3435   0.00887   0.00266  -0.0290   0.7267   1.0000
   1.750   0.3684   0.00899   0.00273  -0.0282   0.7141   1.0000
   2.000   0.3934   0.00910   0.00281  -0.0275   0.7011   1.0000
   2.250   0.4185   0.00922   0.00295  -0.0268   0.6874   1.0000
   2.500   0.4436   0.00932   0.00304  -0.0261   0.6729   1.0000
   2.750   0.4688   0.00943   0.00314  -0.0254   0.6576   1.0000
   3.000   0.4939   0.00953   0.00324  -0.0246   0.6415   1.0000
   3.250   0.5190   0.00964   0.00333  -0.0239   0.6248   1.0000
   3.500   0.5443   0.00973   0.00354  -0.0232   0.6046   1.0000
   3.750   0.5694   0.00983   0.00364  -0.0224   0.5843   1.0000
   4.000   0.5947   0.00992   0.00378  -0.0217   0.5601   1.0000
   4.250   0.6197   0.01005   0.00392  -0.0210   0.5333   1.0000
   4.500   0.6446   0.01022   0.00409  -0.0203   0.5028   1.0000
   4.750   0.6690   0.01047   0.00430  -0.0195   0.4676   1.0000
   5.000   0.6920   0.01086   0.00453  -0.0185   0.4133   1.0000
   5.250   0.7109   0.01182   0.00483  -0.0174   0.2803   1.0000
   5.500   0.7200   0.01507   0.00641  -0.0158   0.0202   1.0000
   5.750   0.7409   0.01658   0.00827  -0.0143   0.0126   1.0000
   6.000   0.7615   0.01790   0.00981  -0.0130   0.0113   1.0000
   6.250   0.7809   0.01946   0.01151  -0.0115   0.0108   1.0000
   6.500   0.8003   0.02127   0.01346  -0.0100   0.0106   1.0000
   6.750   0.8207   0.02335   0.01570  -0.0085   0.0107   1.0000
   7.000   0.8418   0.02579   0.01837  -0.0071   0.0110   1.0000
   7.250   0.8621   0.02862   0.02153  -0.0056   0.0116   1.0000
<< Back to MH 22 7.2% (mh22-il)

Polar data table (+)

Polar graphs


<< Back to MH 22 7.2% (mh22-il)