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

AH-6-40-7 AIRFOIL (ah6407-il) Xfoil prediction polar at RE=1,000,000 Ncrit=5


Details Polar file
Airfoil: AH-6-40-7 AIRFOIL (ah6407-il)
Reynolds number: 1,000,000
Max Cl/Cd: 121.75 at α=4.25°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-ah6407-il-1000000-n5.txt
Download as CSV file: xf-ah6407-il-1000000-n5.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 =     1.000 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -10.500  -0.3309   0.12250   0.12081  -0.0273   1.0000   0.0046
 -10.250  -0.3262   0.11942   0.11775  -0.0281   1.0000   0.0043
 -10.000  -0.3228   0.11621   0.11454  -0.0286   1.0000   0.0040
  -9.750  -0.3214   0.11284   0.11119  -0.0289   1.0000   0.0038
  -8.750  -0.2813   0.09746   0.09583  -0.0397   0.9932   0.0038
  -7.750  -0.2047   0.01795   0.01497  -0.1412   0.9538   0.0051
  -7.500  -0.1734   0.01537   0.01200  -0.1438   0.9499   0.0052
  -7.250  -0.1423   0.01335   0.00962  -0.1457   0.9454   0.0056
  -7.000  -0.1129   0.01236   0.00845  -0.1465   0.9405   0.0060
  -6.750  -0.0838   0.01166   0.00760  -0.1470   0.9358   0.0063
  -6.500  -0.0547   0.01106   0.00686  -0.1474   0.9303   0.0067
  -6.000   0.0034   0.01006   0.00559  -0.1481   0.9196   0.0075
  -5.750   0.0325   0.00958   0.00496  -0.1484   0.9139   0.0084
  -5.500   0.0616   0.00920   0.00450  -0.1487   0.9082   0.0095
  -5.250   0.0905   0.00890   0.00412  -0.1489   0.9018   0.0107
  -5.000   0.1194   0.00862   0.00374  -0.1490   0.8957   0.0123
  -4.750   0.1479   0.00846   0.00357  -0.1491   0.8867   0.0147
  -4.500   0.1761   0.00833   0.00337  -0.1490   0.8735   0.0166
  -4.250   0.2036   0.00838   0.00340  -0.1488   0.8565   0.0193
  -4.000   0.2314   0.00843   0.00338  -0.1486   0.8408   0.0216
  -3.750   0.2597   0.00842   0.00329  -0.1486   0.8285   0.0232
  -3.500   0.2879   0.00846   0.00329  -0.1485   0.8171   0.0241
  -3.250   0.3160   0.00854   0.00331  -0.1484   0.8048   0.0245
  -3.000   0.3454   0.00815   0.00279  -0.1488   0.7913   0.0263
  -2.750   0.3740   0.00801   0.00255  -0.1489   0.7745   0.0276
  -2.500   0.4024   0.00796   0.00239  -0.1489   0.7544   0.0286
  -2.250   0.4308   0.00790   0.00221  -0.1490   0.7337   0.0291
  -2.000   0.4594   0.00783   0.00203  -0.1490   0.7153   0.0294
  -1.750   0.4879   0.00780   0.00190  -0.1491   0.6974   0.0301
  -1.500   0.5162   0.00780   0.00179  -0.1492   0.6759   0.0307
  -1.250   0.5444   0.00782   0.00168  -0.1492   0.6517   0.0312
  -1.000   0.5726   0.00786   0.00161  -0.1492   0.6287   0.0321
  -0.750   0.6006   0.00794   0.00156  -0.1492   0.6046   0.0335
  -0.500   0.6285   0.00803   0.00154  -0.1492   0.5790   0.0348
  -0.250   0.6563   0.00815   0.00154  -0.1492   0.5512   0.0377
   0.000   0.6839   0.00829   0.00156  -0.1492   0.5202   0.0440
   0.250   0.7114   0.00845   0.00161  -0.1491   0.4878   0.0545
   0.500   0.7392   0.00853   0.00169  -0.1493   0.4596   0.1042
   0.750   0.7669   0.00864   0.00177  -0.1493   0.4376   0.1307
   1.000   0.7948   0.00873   0.00186  -0.1494   0.4181   0.1667
   1.250   0.8225   0.00884   0.00195  -0.1494   0.4009   0.1938
   1.500   0.8502   0.00893   0.00205  -0.1495   0.3855   0.2233
   1.750   0.8782   0.00896   0.00219  -0.1496   0.3714   0.2921
   2.000   0.9063   0.00894   0.00234  -0.1499   0.3585   0.3978
   2.250   0.9349   0.00874   0.00255  -0.1503   0.3459   0.6229
   2.500   0.9566   0.00819   0.00265  -0.1490   0.3343   1.0000
   2.750   0.9840   0.00836   0.00277  -0.1489   0.3225   1.0000
   3.000   1.0115   0.00852   0.00288  -0.1489   0.3135   1.0000
   3.250   1.0387   0.00869   0.00302  -0.1488   0.3048   1.0000
   3.500   1.0661   0.00885   0.00315  -0.1487   0.2967   1.0000
   3.750   1.0932   0.00903   0.00329  -0.1486   0.2883   1.0000
   4.250   1.1469   0.00942   0.00362  -0.1483   0.2678   1.0000
   4.500   1.1735   0.00964   0.00380  -0.1482   0.2570   1.0000
   4.750   1.1993   0.00997   0.00403  -0.1479   0.2357   1.0000
   5.000   1.2248   0.01032   0.00427  -0.1476   0.2152   1.0000
   5.250   1.2505   0.01062   0.00450  -0.1473   0.1984   1.0000
   5.500   1.2753   0.01104   0.00479  -0.1468   0.1741   1.0000
   5.750   1.2969   0.01187   0.00531  -0.1460   0.1205   1.0000
   6.000   1.3174   0.01285   0.00600  -0.1449   0.0688   1.0000
   6.250   1.3388   0.01367   0.00662  -0.1439   0.0347   1.0000
   6.500   1.3611   0.01436   0.00719  -0.1431   0.0140   1.0000
   6.750   1.3851   0.01480   0.00763  -0.1424   0.0102   1.0000
   7.000   1.4092   0.01518   0.00804  -0.1419   0.0089   1.0000
   7.250   1.4325   0.01565   0.00854  -0.1411   0.0074   1.0000
   7.500   1.4558   0.01609   0.00902  -0.1404   0.0064   1.0000
   7.750   1.4790   0.01653   0.00949  -0.1397   0.0058   1.0000
   8.000   1.5015   0.01702   0.01001  -0.1389   0.0052   1.0000
   8.250   1.5227   0.01763   0.01066  -0.1378   0.0046   1.0000
   8.500   1.5442   0.01817   0.01126  -0.1368   0.0044   1.0000
   8.750   1.5654   0.01871   0.01187  -0.1358   0.0041   1.0000
   9.000   1.5860   0.01927   0.01248  -0.1347   0.0038   1.0000
   9.250   1.6061   0.01984   0.01309  -0.1335   0.0035   1.0000
   9.500   1.6250   0.02048   0.01378  -0.1321   0.0032   1.0000
   9.750   1.6409   0.02136   0.01475  -0.1302   0.0029   1.0000
  10.000   1.6572   0.02212   0.01558  -0.1284   0.0028   1.0000
  10.250   1.6726   0.02286   0.01640  -0.1265   0.0027   1.0000
  10.500   1.6861   0.02367   0.01729  -0.1242   0.0026   1.0000
  10.750   1.6978   0.02452   0.01823  -0.1217   0.0025   1.0000
  11.000   1.7059   0.02542   0.01925  -0.1185   0.0024   1.0000
  11.250   1.7129   0.02645   0.02036  -0.1153   0.0023   1.0000
  11.500   1.7200   0.02753   0.02154  -0.1123   0.0023   1.0000
  11.750   1.7265   0.02872   0.02282  -0.1094   0.0022   1.0000
  12.000   1.7323   0.03001   0.02420  -0.1066   0.0021   1.0000
  12.250   1.7376   0.03140   0.02569  -0.1040   0.0021   1.0000
  12.500   1.7416   0.03296   0.02735  -0.1014   0.0020   1.0000
  12.750   1.7446   0.03468   0.02917  -0.0989   0.0019   1.0000
  13.000   1.7449   0.03674   0.03134  -0.0965   0.0019   1.0000
  13.250   1.7431   0.03911   0.03383  -0.0941   0.0018   1.0000
  13.500   1.7377   0.04197   0.03683  -0.0919   0.0018   1.0000
  13.750   1.7293   0.04537   0.04037  -0.0900   0.0017   1.0000
  14.000   1.7209   0.04896   0.04411  -0.0885   0.0017   1.0000
  14.250   1.7167   0.05229   0.04757  -0.0877   0.0017   1.0000
  14.500   1.7108   0.05600   0.05142  -0.0871   0.0017   1.0000
  14.750   1.7041   0.06001   0.05558  -0.0870   0.0017   1.0000
  15.000   1.6946   0.06465   0.06036  -0.0872   0.0016   1.0000
  15.250   1.6845   0.06966   0.06552  -0.0880   0.0016   1.0000
  15.500   1.6746   0.07485   0.07085  -0.0891   0.0016   1.0000
  15.750   1.6599   0.08101   0.07716  -0.0907   0.0016   1.0000
  16.000   1.6479   0.08692   0.08321  -0.0925   0.0016   1.0000
  16.250   1.6325   0.09361   0.09005  -0.0947   0.0016   1.0000
  16.500   1.6173   0.10035   0.09692  -0.0971   0.0015   1.0000
  16.750   1.5998   0.10756   0.10428  -0.0998   0.0015   1.0000
  17.000   1.5840   0.11459   0.11143  -0.1026   0.0015   1.0000
<< Back to AH-6-40-7 AIRFOIL (ah6407-il)

Polar data table (+)

Polar graphs


<< Back to AH-6-40-7 AIRFOIL (ah6407-il)