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

GOE 417A (GEW. PLATTE) AIRFOIL (goe417a-il) Xfoil prediction polar at RE=100,000 Ncrit=5


Details Polar file
Airfoil: GOE 417A (GEW. PLATTE) AIRFOIL (goe417a-il)
Reynolds number: 100,000
Max Cl/Cd: 56.41 at α=5.25°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-goe417a-il-100000-n5.txt
Download as CSV file: xf-goe417a-il-100000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 417A (GEW. PLATTE) 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
  ------ -------- --------- --------- -------- -------- --------
  -8.250  -0.3462   0.10942   0.10427  -0.0178   1.0000   0.0289
  -8.000  -0.3500   0.10830   0.10321  -0.0167   1.0000   0.0290
  -7.750  -0.3555   0.10741   0.10237  -0.0152   1.0000   0.0291
  -7.500  -0.3570   0.10633   0.10134  -0.0150   1.0000   0.0292
  -7.250  -0.3555   0.10510   0.10013  -0.0155   1.0000   0.0293
  -7.000  -0.3517   0.10386   0.09892  -0.0167   1.0000   0.0294
  -6.500  -0.3428   0.09787   0.09300  -0.0161   1.0000   0.0297
  -6.250  -0.3409   0.09380   0.08897  -0.0138   1.0000   0.0300
  -6.000  -0.3369   0.09092   0.08612  -0.0130   1.0000   0.0304
  -5.750  -0.3311   0.08842   0.08364  -0.0129   1.0000   0.0307
  -5.500  -0.3240   0.08600   0.08124  -0.0132   1.0000   0.0311
  -5.250  -0.3157   0.08359   0.07884  -0.0136   1.0000   0.0315
  -5.000  -0.3019   0.08097   0.07621  -0.0153   0.9991   0.0319
  -4.750  -0.2712   0.07759   0.07277  -0.0209   0.9948   0.0327
  -4.500  -0.2389   0.07432   0.06943  -0.0266   0.9897   0.0337
  -4.250  -0.1976   0.07136   0.06638  -0.0342   0.9846   0.0352
  -4.000  -0.1381   0.07029   0.06505  -0.0457   0.9774   0.0363
  -3.750  -0.0903   0.06833   0.06289  -0.0535   0.9712   0.0365
  -3.500  -0.0692   0.06344   0.05804  -0.0553   0.9661   0.0368
  -3.250  -0.0509   0.05931   0.05395  -0.0562   0.9610   0.0374
  -3.000  -0.0254   0.05621   0.05082  -0.0585   0.9556   0.0381
  -2.750   0.0063   0.05350   0.04804  -0.0619   0.9511   0.0391
  -2.500   0.0351   0.05119   0.04566  -0.0644   0.9443   0.0406
  -2.250   0.0744   0.04894   0.04330  -0.0687   0.9402   0.0433
  -2.000   0.1150   0.04824   0.04237  -0.0721   0.9323   0.0453
  -1.750   0.1638   0.04763   0.04146  -0.0767   0.9277   0.0458
  -1.500   0.1800   0.04388   0.03780  -0.0766   0.9197   0.0464
  -1.250   0.2056   0.04092   0.03486  -0.0780   0.9143   0.0476
  -1.000   0.2315   0.03901   0.03290  -0.0787   0.9066   0.0493
  -0.750   0.2642   0.03729   0.03109  -0.0804   0.9000   0.0517
  -0.500   0.3020   0.03660   0.03019  -0.0820   0.8929   0.0553
  -0.250   0.3348   0.03522   0.02864  -0.0829   0.8850   0.0568
   0.000   0.3591   0.03268   0.02617  -0.0835   0.8783   0.0596
   0.250   0.3879   0.03156   0.02495  -0.0837   0.8689   0.0649
   0.750   0.4485   0.02868   0.02188  -0.0845   0.8514   0.0713
   1.000   0.4759   0.02763   0.02072  -0.0840   0.8384   0.0771
   1.250   0.5067   0.02653   0.01946  -0.0839   0.8247   0.0819
   1.500   0.5357   0.02508   0.01797  -0.0839   0.8097   0.0882
   1.750   0.5691   0.02389   0.01665  -0.0845   0.7944   0.0966
   2.000   0.6052   0.02271   0.01535  -0.0857   0.7790   0.1090
   2.250   0.6421   0.02149   0.01402  -0.0873   0.7630   0.1351
   2.500   0.6795   0.02027   0.01272  -0.0891   0.7448   0.1684
   2.750   0.7205   0.01946   0.01170  -0.0908   0.7242   0.1713
   3.500   0.8278   0.01849   0.00985  -0.0910   0.6538   0.0542
   3.750   0.8570   0.01815   0.00941  -0.0908   0.6280   0.0527
   4.000   0.8879   0.01794   0.00910  -0.0911   0.6044   0.0529
   4.250   0.9215   0.01789   0.00892  -0.0920   0.5804   0.0552
   4.500   0.9573   0.01780   0.00874  -0.0934   0.5541   0.0555
   4.750   0.9886   0.01779   0.00860  -0.0939   0.5222   0.0558
   5.000   1.0122   0.01796   0.00858  -0.0928   0.4792   0.0565
   5.250   1.0306   0.01827   0.00865  -0.0908   0.4167   0.0578
   5.500   1.0451   0.01891   0.00887  -0.0882   0.3346   0.0622
   5.750   1.0606   0.01965   0.00922  -0.0859   0.2817   0.0669
   6.000   1.0771   0.02034   0.00967  -0.0839   0.2477   0.0704
   6.250   1.0938   0.02102   0.01017  -0.0819   0.2201   0.0766
   6.500   1.1112   0.02162   0.01069  -0.0799   0.1984   0.0890
   6.750   1.1843   0.02284   0.01219  -0.0917   0.0772   1.0000
   7.000   1.1960   0.02413   0.01325  -0.0888   0.0611   1.0000
   7.250   1.2082   0.02537   0.01442  -0.0860   0.0534   1.0000
   7.500   1.2225   0.02635   0.01550  -0.0835   0.0484   1.0000
   7.750   1.2348   0.02749   0.01674  -0.0808   0.0453   1.0000
   8.000   1.2432   0.02894   0.01825  -0.0775   0.0429   1.0000
   8.250   1.2551   0.03006   0.01948  -0.0747   0.0408   1.0000
   8.500   1.2666   0.03122   0.02075  -0.0720   0.0381   1.0000
   8.750   1.2768   0.03249   0.02210  -0.0692   0.0360   1.0000
   9.000   1.2856   0.03394   0.02361  -0.0662   0.0347   1.0000
   9.250   1.2952   0.03565   0.02533  -0.0635   0.0335   1.0000
   9.500   1.3105   0.03781   0.02750  -0.0618   0.0323   1.0000
   9.750   1.3240   0.03917   0.02908  -0.0595   0.0310   1.0000
  10.000   1.3365   0.04071   0.03080  -0.0572   0.0293   1.0000
  10.250   1.3485   0.04243   0.03270  -0.0550   0.0280   1.0000
  10.500   1.3612   0.04444   0.03488  -0.0531   0.0271   1.0000
  10.750   1.3721   0.04657   0.03719  -0.0509   0.0264   1.0000
  11.000   1.3813   0.04889   0.03966  -0.0487   0.0258   1.0000
  11.250   1.3895   0.05154   0.04249  -0.0466   0.0253   1.0000
  11.500   1.3958   0.05484   0.04596  -0.0445   0.0248   1.0000
  11.750   1.3893   0.05728   0.04872  -0.0406   0.0245   1.0000
  12.000   1.3794   0.05961   0.05138  -0.0366   0.0242   1.0000
  12.250   1.3676   0.06226   0.05436  -0.0331   0.0238   1.0000
  12.500   1.3532   0.06530   0.05773  -0.0301   0.0234   1.0000
  12.750   1.3392   0.06875   0.06145  -0.0278   0.0232   1.0000
  13.000   1.3232   0.07253   0.06550  -0.0263   0.0230   1.0000
  13.250   1.3055   0.07684   0.07007  -0.0256   0.0229   1.0000
  13.500   1.2873   0.08171   0.07518  -0.0257   0.0230   1.0000
  13.750   1.2670   0.08712   0.08083  -0.0270   0.0228   1.0000
  14.000   1.2465   0.09323   0.08715  -0.0291   0.0230   1.0000
  14.250   1.2231   0.10035   0.09448  -0.0325   0.0230   1.0000
<< Back to GOE 417A (GEW. PLATTE) AIRFOIL (goe417a-il)

Polar data table (+)

Polar graphs


<< Back to GOE 417A (GEW. PLATTE) AIRFOIL (goe417a-il)