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

AH-7-47-6 AIRFOIL (ah7476-il) Xfoil prediction polar at RE=100,000 Ncrit=5


Details Polar file
Airfoil: AH-7-47-6 AIRFOIL (ah7476-il)
Reynolds number: 100,000
Max Cl/Cd: 74.44 at α=4.25°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-ah7476-il-100000-n5.txt
Download as CSV file: xf-ah7476-il-100000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: AH-7-47-6 AIRFOIL                               
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.100 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.000  -0.3438   0.11577   0.11104  -0.0218   1.0000   0.0294
  -7.750  -0.3536   0.11491   0.11027  -0.0199   1.0000   0.0295
  -7.500  -0.3637   0.11403   0.10946  -0.0180   1.0000   0.0295
  -7.250  -0.3498   0.11110   0.10655  -0.0242   0.9962   0.0297
  -7.000  -0.3389   0.10575   0.10121  -0.0221   0.9952   0.0303
  -6.750  -0.3253   0.10202   0.09745  -0.0232   0.9924   0.0317
  -6.500  -0.3082   0.09855   0.09398  -0.0271   0.9886   0.0336
  -6.250  -0.2874   0.09496   0.09038  -0.0330   0.9838   0.0356
  -6.000  -0.2454   0.09121   0.08658  -0.0499   0.9779   0.0370
  -5.750  -0.2345   0.08695   0.08234  -0.0500   0.9740   0.0376
  -5.500  -0.2223   0.08338   0.07877  -0.0491   0.9712   0.0391
  -5.250  -0.1971   0.07984   0.07520  -0.0540   0.9682   0.0422
  -5.000  -0.1546   0.07528   0.07057  -0.0676   0.9626   0.0469
  -4.750  -0.1385   0.07218   0.06747  -0.0674   0.9600   0.0497
  -4.500  -0.0722   0.06698   0.06207  -0.0869   0.9560   0.0567
  -4.250  -0.0648   0.06420   0.05934  -0.0837   0.9522   0.0588
  -4.000  -0.0045   0.05958   0.05453  -0.0982   0.9500   0.0687
  -3.750   0.0179   0.05687   0.05182  -0.0989   0.9478   0.0737
  -3.500   0.0731   0.05253   0.04727  -0.1110   0.9438   0.0820
  -3.250   0.1374   0.04882   0.04322  -0.1238   0.9417   0.0941
  -3.000   0.1569   0.04623   0.04074  -0.1232   0.9388   0.0986
  -2.500   0.3133   0.03434   0.02753  -0.1504   0.9378   0.0572
  -2.250   0.3648   0.03171   0.02435  -0.1562   0.9348   0.0559
  -2.000   0.4044   0.02963   0.02208  -0.1596   0.9299   0.0544
  -1.750   0.4517   0.02757   0.01963  -0.1641   0.9271   0.0521
  -1.500   0.4997   0.02580   0.01741  -0.1683   0.9248   0.0504
  -1.250   0.5343   0.02469   0.01597  -0.1697   0.9180   0.0494
  -1.000   0.5758   0.02355   0.01455  -0.1723   0.9137   0.0490
  -0.500   0.6439   0.02225   0.01305  -0.1747   0.9028   0.0529
  -0.250   0.6815   0.02156   0.01230  -0.1765   0.8993   0.0530
   0.000   0.7214   0.02088   0.01160  -0.1788   0.8967   0.0532
   0.250   0.7470   0.02068   0.01141  -0.1783   0.8879   0.0536
   0.500   0.7851   0.02017   0.01088  -0.1802   0.8839   0.0544
   0.750   0.8158   0.01992   0.01061  -0.1806   0.8768   0.0554
   1.000   0.8500   0.01959   0.01026  -0.1816   0.8706   0.0574
   1.250   0.8885   0.01914   0.00982  -0.1833   0.8659   0.0621
   1.500   0.9167   0.01898   0.00974  -0.1831   0.8559   0.0786
   1.750   0.9486   0.01715   0.00968  -0.1839   0.8497   1.0000
   2.250   1.0087   0.01699   0.00942  -0.1839   0.8284   1.0000
   2.500   1.0424   0.01680   0.00918  -0.1846   0.8176   1.0000
   2.750   1.0765   0.01660   0.00897  -0.1854   0.8054   1.0000
   3.000   1.1041   0.01656   0.00895  -0.1848   0.7855   1.0000
   3.250   1.1341   0.01639   0.00872  -0.1845   0.7582   1.0000
   3.500   1.1600   0.01635   0.00863  -0.1834   0.7206   1.0000
   3.750   1.1899   0.01633   0.00850  -0.1831   0.6803   1.0000
   4.000   1.2198   0.01644   0.00845  -0.1830   0.6418   1.0000
   4.250   1.2468   0.01675   0.00858  -0.1824   0.6047   1.0000
   4.500   1.2714   0.01719   0.00888  -0.1814   0.5709   1.0000
   4.750   1.2944   0.01773   0.00930  -0.1803   0.5376   1.0000
   5.000   1.3161   0.01832   0.00976  -0.1789   0.5057   1.0000
   5.250   1.3369   0.01896   0.01027  -0.1774   0.4754   1.0000
   5.500   1.3574   0.01961   0.01083  -0.1759   0.4477   1.0000
   5.750   1.3773   0.02028   0.01147  -0.1743   0.4213   1.0000
   6.000   1.3962   0.02096   0.01211  -0.1726   0.3924   1.0000
   6.250   1.4150   0.02160   0.01278  -0.1708   0.3621   1.0000
   6.500   1.4325   0.02229   0.01345  -0.1689   0.3245   1.0000
   6.750   1.4459   0.02326   0.01419  -0.1663   0.2733   1.0000
   7.250   1.4701   0.02567   0.01618  -0.1611   0.1906   1.0000
   7.500   1.4820   0.02696   0.01729  -0.1586   0.1541   1.0000
   7.750   1.4940   0.02822   0.01842  -0.1561   0.1184   1.0000
   8.250   1.5088   0.03134   0.02122  -0.1496   0.0596   1.0000
   8.500   1.5135   0.03302   0.02291  -0.1458   0.0446   1.0000
   8.750   1.5193   0.03466   0.02460  -0.1423   0.0347   1.0000
   9.000   1.5258   0.03625   0.02629  -0.1390   0.0294   1.0000
   9.250   1.5306   0.03798   0.02812  -0.1357   0.0267   1.0000
   9.500   1.5334   0.03989   0.03018  -0.1321   0.0251   1.0000
   9.750   1.5360   0.04184   0.03234  -0.1287   0.0238   1.0000
  10.000   1.5381   0.04388   0.03458  -0.1254   0.0225   1.0000
  10.250   1.5398   0.04602   0.03691  -0.1224   0.0211   1.0000
  10.500   1.5397   0.04837   0.03942  -0.1194   0.0199   1.0000
  10.750   1.5363   0.05115   0.04233  -0.1164   0.0190   1.0000
  11.000   1.5325   0.05420   0.04552  -0.1134   0.0184   1.0000
  11.250   1.5335   0.05690   0.04845  -0.1109   0.0181   1.0000
  11.500   1.5346   0.05974   0.05152  -0.1086   0.0177   1.0000
  11.750   1.5355   0.06274   0.05477  -0.1064   0.0174   1.0000
  12.000   1.5357   0.06594   0.05822  -0.1044   0.0171   1.0000
  12.250   1.5343   0.06939   0.06195  -0.1025   0.0168   1.0000
  12.500   1.5306   0.07310   0.06594  -0.1008   0.0164   1.0000
  12.750   1.5247   0.07709   0.07024  -0.0995   0.0161   1.0000
  13.000   1.5165   0.08139   0.07482  -0.0986   0.0158   1.0000
  13.250   1.5063   0.08600   0.07971  -0.0982   0.0154   1.0000
  13.500   1.4946   0.09099   0.08497  -0.0984   0.0152   1.0000
  13.750   1.4816   0.09644   0.09069  -0.0993   0.0150   1.0000
  14.000   1.4676   0.10236   0.09687  -0.1009   0.0148   1.0000
  14.250   1.4527   0.10886   0.10363  -0.1034   0.0147   1.0000
  14.500   1.4370   0.11600   0.11103  -0.1068   0.0147   1.0000
  14.750   1.4204   0.12388   0.11916  -0.1112   0.0147   1.0000
  15.000   1.4025   0.13270   0.12821  -0.1167   0.0149   1.0000
  15.250   1.3823   0.14290   0.13865  -0.1236   0.0152   1.0000
  15.500   1.3601   0.15473   0.15067  -0.1320   0.0156   1.0000
  15.750   1.3369   0.16835   0.16446  -0.1418   0.0163   1.0000
<< Back to AH-7-47-6 AIRFOIL (ah7476-il)

Polar data table (+)

Polar graphs


<< Back to AH-7-47-6 AIRFOIL (ah7476-il)