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

GOE 602 MOD. AIRFOIL (goe602m-il) Xfoil prediction polar at RE=50,000 Ncrit=5


Details Polar file
Airfoil: GOE 602 MOD. AIRFOIL (goe602m-il)
Reynolds number: 50,000
Max Cl/Cd: 40.52 at α=6.75°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-goe602m-il-50000-n5.txt
Download as CSV file: xf-goe602m-il-50000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 602 MOD. 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 =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.500  -0.3651   0.09866   0.09193  -0.0382   1.0000   0.0586
  -8.250  -0.3695   0.09575   0.08912  -0.0379   1.0000   0.0581
  -8.000  -0.3764   0.09304   0.08652  -0.0372   1.0000   0.0579
  -7.750  -0.3856   0.09051   0.08410  -0.0362   1.0000   0.0575
  -7.500  -0.3954   0.08788   0.08160  -0.0354   1.0000   0.0572
  -7.250  -0.4028   0.08473   0.07855  -0.0355   1.0000   0.0569
  -7.000  -0.4088   0.08139   0.07530  -0.0362   1.0000   0.0565
  -6.750  -0.4135   0.07765   0.07161  -0.0375   1.0000   0.0562
  -6.500  -0.4158   0.07346   0.06746  -0.0394   1.0000   0.0557
  -6.250  -0.4156   0.06872   0.06270  -0.0419   1.0000   0.0551
  -6.000  -0.4123   0.06345   0.05734  -0.0449   1.0000   0.0544
  -5.750  -0.4052   0.05799   0.05168  -0.0481   1.0000   0.0539
  -5.500  -0.3935   0.05256   0.04593  -0.0513   1.0000   0.0537
  -5.250  -0.3778   0.04791   0.04088  -0.0537   1.0000   0.0554
  -5.000  -0.3582   0.04349   0.03586  -0.0560   1.0000   0.0583
  -4.750  -0.3351   0.03932   0.03092  -0.0576   1.0000   0.0604
  -4.500  -0.3117   0.03607   0.02706  -0.0583   1.0000   0.0620
  -4.250  -0.2917   0.03457   0.02549  -0.0582   1.0000   0.0667
  -4.000  -0.2627   0.03251   0.02285  -0.0592   0.9981   0.0722
  -3.750  -0.2253   0.03039   0.02011  -0.0612   0.9932   0.0753
  -3.500  -0.1899   0.02925   0.01878  -0.0634   0.9873   0.0832
  -3.250  -0.1543   0.02805   0.01721  -0.0650   0.9810   0.0899
  -3.000  -0.1181   0.02704   0.01604  -0.0668   0.9750   0.0955
  -2.750  -0.0830   0.02624   0.01492  -0.0681   0.9679   0.1032
  -2.500  -0.0475   0.02547   0.01401  -0.0699   0.9607   0.1164
  -2.250  -0.0090   0.02455   0.01320  -0.0725   0.9541   0.1519
  -2.000   0.0257   0.02364   0.01294  -0.0745   0.9461   0.2875
  -1.750   0.0613   0.02313   0.01306  -0.0763   0.9390   0.4501
  -1.500   0.0896   0.02283   0.01305  -0.0762   0.9290   0.5634
  -1.250   0.1202   0.02236   0.01286  -0.0763   0.9205   0.6592
  -1.000   0.1625   0.02150   0.01248  -0.0781   0.9121   1.0000
  -0.750   0.1971   0.02167   0.01226  -0.0798   0.9014   1.0000
  -0.500   0.2345   0.02182   0.01206  -0.0819   0.8919   1.0000
  -0.250   0.2722   0.02194   0.01190  -0.0839   0.8825   1.0000
   0.000   0.3038   0.02208   0.01182  -0.0847   0.8711   1.0000
   0.250   0.3369   0.02220   0.01175  -0.0858   0.8603   1.0000
   0.500   0.3747   0.02224   0.01162  -0.0876   0.8514   1.0000
   0.750   0.4080   0.02232   0.01157  -0.0885   0.8405   1.0000
   1.250   0.4685   0.02252   0.01159  -0.0892   0.8166   1.0000
   1.500   0.5003   0.02257   0.01157  -0.0897   0.8053   1.0000
   1.750   0.5358   0.02251   0.01146  -0.0907   0.7954   1.0000
   2.000   0.5660   0.02256   0.01149  -0.0909   0.7830   1.0000
   2.250   0.5951   0.02263   0.01155  -0.0908   0.7700   1.0000
   2.500   0.6245   0.02269   0.01161  -0.0908   0.7570   1.0000
   2.750   0.6543   0.02274   0.01169  -0.0908   0.7438   1.0000
   3.000   0.6843   0.02279   0.01176  -0.0908   0.7305   1.0000
   3.250   0.7143   0.02283   0.01184  -0.0907   0.7168   1.0000
   3.500   0.7439   0.02290   0.01197  -0.0906   0.7026   1.0000
   3.750   0.7729   0.02299   0.01211  -0.0903   0.6879   1.0000
   4.000   0.8014   0.02311   0.01229  -0.0900   0.6727   1.0000
   4.250   0.8294   0.02327   0.01255  -0.0896   0.6571   1.0000
   4.500   0.8570   0.02345   0.01281  -0.0891   0.6413   1.0000
   4.750   0.8843   0.02366   0.01311  -0.0886   0.6253   1.0000
   5.000   0.9114   0.02391   0.01348  -0.0880   0.6093   1.0000
   5.250   0.9381   0.02419   0.01387  -0.0874   0.5933   1.0000
   5.500   0.9628   0.02457   0.01438  -0.0866   0.5766   1.0000
   5.750   0.9862   0.02500   0.01496  -0.0856   0.5591   1.0000
   6.000   1.0095   0.02539   0.01554  -0.0845   0.5406   1.0000
   6.250   1.0337   0.02571   0.01596  -0.0833   0.5210   1.0000
   6.500   1.0530   0.02614   0.01653  -0.0815   0.4983   1.0000
   6.750   1.0733   0.02649   0.01696  -0.0797   0.4746   1.0000
   7.000   1.0838   0.02685   0.01730  -0.0763   0.4381   1.0000
   7.250   1.0900   0.02733   0.01762  -0.0724   0.3932   1.0000
   7.500   1.1006   0.02801   0.01834  -0.0696   0.3586   1.0000
   7.750   1.1098   0.02884   0.01918  -0.0667   0.3215   1.0000
   8.000   1.1161   0.02986   0.02009  -0.0636   0.2775   1.0000
   8.250   1.1182   0.03120   0.02119  -0.0601   0.2263   1.0000
   8.500   1.1193   0.03294   0.02264  -0.0569   0.1685   1.0000
   8.750   1.1140   0.03547   0.02463  -0.0537   0.1069   1.0000
   9.000   1.1118   0.03803   0.02693  -0.0511   0.0726   1.0000
   9.250   1.1124   0.04040   0.02924  -0.0488   0.0537   1.0000
   9.500   1.1125   0.04285   0.03169  -0.0468   0.0470   1.0000
   9.750   1.1112   0.04548   0.03433  -0.0450   0.0439   1.0000
  10.000   1.1107   0.04811   0.03715  -0.0434   0.0417   1.0000
  10.250   1.1096   0.05091   0.04013  -0.0421   0.0399   1.0000
  10.500   1.1075   0.05394   0.04330  -0.0411   0.0382   1.0000
  10.750   1.1045   0.05718   0.04667  -0.0404   0.0367   1.0000
  11.000   1.1004   0.06066   0.05030  -0.0399   0.0354   1.0000
  11.250   1.1017   0.06368   0.05357  -0.0393   0.0340   1.0000
  11.500   1.1036   0.06670   0.05681  -0.0387   0.0328   1.0000
  11.750   1.1074   0.06960   0.05992  -0.0380   0.0320   1.0000
  12.000   1.1133   0.07236   0.06288  -0.0372   0.0312   1.0000
  12.250   1.1211   0.07506   0.06581  -0.0363   0.0306   1.0000
  12.500   1.1293   0.07790   0.06889  -0.0354   0.0299   1.0000
  12.750   1.1348   0.08117   0.07239  -0.0350   0.0292   1.0000
  13.000   1.1370   0.08488   0.07631  -0.0350   0.0286   1.0000
  13.250   1.1372   0.08895   0.08055  -0.0352   0.0279   1.0000
  13.500   1.1353   0.09354   0.08533  -0.0358   0.0274   1.0000
  13.750   1.1266   0.09888   0.09088  -0.0376   0.0272   1.0000
  14.000   1.1155   0.10467   0.09694  -0.0401   0.0272   1.0000
  14.250   1.1033   0.11094   0.10344  -0.0432   0.0272   1.0000
  14.500   1.0910   0.11758   0.11028  -0.0467   0.0272   1.0000
  14.750   1.0783   0.12466   0.11755  -0.0507   0.0273   1.0000
  15.000   1.0600   0.13373   0.12694  -0.0569   0.0279   1.0000
  15.250   1.0240   0.15035   0.14390  -0.0681   0.0301   1.0000
<< Back to GOE 602 MOD. AIRFOIL (goe602m-il)

Polar data table (+)

Polar graphs


<< Back to GOE 602 MOD. AIRFOIL (goe602m-il)