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AH-6-40-7 AIRFOIL (ah6407-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: AH-6-40-7 AIRFOIL (ah6407-il)
Reynolds number: 50,000
Max Cl/Cd: 41.61 at α=5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-ah6407-il-50000.txt
Download as CSV file: xf-ah6407-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: AH-6-40-7 AIRFOIL                               
  
 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
  ------ -------- --------- --------- -------- -------- --------
  -8.000  -0.3241   0.10991   0.10310  -0.0245   1.0000   0.1324
  -7.750  -0.3334   0.10947   0.10279  -0.0239   1.0000   0.1358
  -7.500  -0.3469   0.11018   0.10368  -0.0248   1.0000   0.1371
  -7.250  -0.3306   0.10369   0.09718  -0.0215   1.0000   0.1431
  -7.000  -0.3330   0.10208   0.09568  -0.0211   1.0000   0.1488
  -6.750  -0.3421   0.10274   0.09648  -0.0246   1.0000   0.1517
  -6.500  -0.3325   0.09717   0.09094  -0.0201   1.0000   0.1569
  -6.250  -0.3317   0.09524   0.08909  -0.0202   1.0000   0.1633
  -6.000  -0.3317   0.09433   0.08827  -0.0249   1.0000   0.1673
  -5.750  -0.3263   0.08994   0.08392  -0.0199   1.0000   0.1727
  -5.500  -0.3171   0.08966   0.08368  -0.0289   1.0000   0.1813
  -5.250  -0.3159   0.08481   0.07891  -0.0217   1.0000   0.1857
  -5.000  -0.3017   0.08307   0.07717  -0.0285   1.0000   0.1964
  -4.750  -0.2992   0.07926   0.07341  -0.0233   1.0000   0.2024
  -4.500  -0.2865   0.07643   0.07060  -0.0260   1.0000   0.2138
  -4.250  -0.2732   0.07356   0.06774  -0.0278   1.0000   0.2278
  -4.000  -0.2600   0.07061   0.06481  -0.0287   1.0000   0.2430
  -3.750  -0.2453   0.06765   0.06186  -0.0296   1.0000   0.2589
  -3.500  -0.2275   0.06480   0.05900  -0.0314   1.0000   0.2785
  -2.250  -0.1580   0.05122   0.04557  -0.0272   1.0000   0.4617
  -1.750   0.1015   0.03862   0.03062  -0.0884   1.0000   0.1983
  -1.500   0.1532   0.03544   0.02675  -0.0944   1.0000   0.1771
  -1.250   0.1933   0.03363   0.02443  -0.0977   1.0000   0.1789
  -1.000   0.2300   0.03209   0.02245  -0.1000   1.0000   0.1781
  -0.750   0.2657   0.03096   0.02081  -0.1020   1.0000   0.1807
  -0.500   0.2950   0.03035   0.02007  -0.1030   1.0000   0.1921
  -0.250   0.3249   0.02996   0.01946  -0.1041   1.0000   0.2091
   0.000   0.3541   0.02959   0.01893  -0.1047   1.0000   0.2295
   0.250   0.3851   0.02919   0.01857  -0.1058   1.0000   0.2859
   0.500   0.4160   0.02636   0.01785  -0.1060   1.0000   0.9296
   0.750   0.4388   0.02734   0.01809  -0.1063   1.0000   1.0000
   1.000   0.4590   0.02842   0.01889  -0.1065   1.0000   1.0000
   1.250   0.4881   0.02960   0.01984  -0.1086   0.9955   1.0000
   1.500   0.5508   0.03050   0.02046  -0.1164   0.9736   1.0000
   1.750   0.6078   0.03115   0.02096  -0.1226   0.9505   1.0000
   2.000   0.6688   0.03147   0.02119  -0.1289   0.9287   1.0000
   2.250   0.7168   0.03160   0.02129  -0.1325   0.9030   1.0000
   2.500   0.7655   0.03148   0.02118  -0.1357   0.8780   1.0000
   2.750   0.8247   0.03072   0.02050  -0.1399   0.8567   1.0000
   3.000   0.8683   0.03014   0.01999  -0.1412   0.8308   1.0000
   3.250   0.9106   0.02939   0.01936  -0.1418   0.8054   1.0000
   3.500   0.9545   0.02839   0.01843  -0.1422   0.7813   1.0000
   3.750   0.9955   0.02745   0.01755  -0.1420   0.7564   1.0000
   4.000   1.0290   0.02698   0.01712  -0.1410   0.7285   1.0000
   4.250   1.0603   0.02678   0.01698  -0.1398   0.7008   1.0000
   4.500   1.0895   0.02688   0.01708  -0.1387   0.6738   1.0000
   4.750   1.1178   0.02718   0.01739  -0.1376   0.6488   1.0000
   5.000   1.1469   0.02756   0.01780  -0.1367   0.6265   1.0000
   5.250   1.1714   0.02841   0.01873  -0.1357   0.6044   1.0000
   5.500   1.1979   0.02919   0.01957  -0.1349   0.5852   1.0000
   5.750   1.2248   0.03003   0.02046  -0.1341   0.5680   1.0000
   6.000   1.2486   0.03115   0.02177  -0.1333   0.5508   1.0000
   6.250   1.2715   0.03236   0.02317  -0.1323   0.5342   1.0000
   6.500   1.2941   0.03351   0.02450  -0.1312   0.5171   1.0000
   6.750   1.3173   0.03456   0.02571  -0.1299   0.4997   1.0000
   7.000   1.3416   0.03549   0.02683  -0.1287   0.4821   1.0000
   7.250   1.3637   0.03653   0.02805  -0.1271   0.4633   1.0000
   7.500   1.3846   0.03716   0.02884  -0.1250   0.4393   1.0000
   7.750   1.4061   0.03681   0.02840  -0.1222   0.4062   1.0000
   8.000   1.4179   0.03583   0.02727  -0.1180   0.3621   1.0000
   8.250   1.4241   0.03544   0.02692  -0.1135   0.3192   1.0000
   8.500   1.4177   0.03576   0.02727  -0.1074   0.2638   1.0000
   8.750   1.4005   0.03762   0.02865  -0.1004   0.1927   1.0000
   9.000   1.3936   0.04053   0.03104  -0.0954   0.1458   1.0000
   9.250   1.3988   0.04361   0.03375  -0.0921   0.1230   1.0000
   9.500   1.4115   0.04663   0.03674  -0.0897   0.1071   1.0000
   9.750   1.4356   0.05007   0.04012  -0.0887   0.0964   1.0000
  10.000   1.4598   0.05363   0.04395  -0.0877   0.0899   1.0000
  10.250   1.4801   0.05783   0.04831  -0.0868   0.0850   1.0000
  10.500   1.4831   0.06140   0.05248  -0.0839   0.0821   1.0000
  10.750   1.4850   0.06525   0.05675  -0.0812   0.0799   1.0000
  11.000   1.4820   0.06941   0.06134  -0.0782   0.0792   1.0000
  11.250   1.4701   0.07350   0.06582  -0.0747   0.0791   1.0000
  11.500   1.4516   0.07767   0.07036  -0.0713   0.0793   1.0000
  11.750   1.4288   0.08224   0.07527  -0.0687   0.0798   1.0000
  12.000   1.4028   0.08740   0.08073  -0.0673   0.0805   1.0000
  12.250   1.3745   0.09328   0.08689  -0.0674   0.0813   1.0000
  12.500   1.3448   0.10007   0.09391  -0.0691   0.0822   1.0000
  12.750   1.3150   0.10786   0.10184  -0.0725   0.0833   1.0000
  13.000   1.2878   0.11649   0.11060  -0.0770   0.0845   1.0000
  13.250   1.2663   0.12530   0.11948  -0.0817   0.0857   1.0000
  13.500   1.2504   0.13399   0.12824  -0.0862   0.0867   1.0000
  13.750   0.9035   0.15771   0.15187  -0.0887   0.1360   1.0000
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