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 AIRFOIL (goe602-il) Xfoil prediction polar at RE=500,000 Ncrit=5


Details Polar file
Airfoil: GOE 602 AIRFOIL (goe602-il)
Reynolds number: 500,000
Max Cl/Cd: 93.19 at α=4.25°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-goe602-il-500000-n5.txt
Download as CSV file: xf-goe602-il-500000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 602 AIRFOIL                                 
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.500 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -10.500  -0.4047   0.10090   0.09862  -0.0367   1.0000   0.0067
 -10.250  -0.4072   0.09736   0.09511  -0.0373   1.0000   0.0068
  -9.750  -0.6002   0.03362   0.03087  -0.0871   0.9798   0.0063
  -9.500  -0.5914   0.02680   0.02333  -0.0907   0.9745   0.0064
  -9.250  -0.5740   0.02345   0.01952  -0.0916   0.9695   0.0066
  -9.000  -0.5485   0.02139   0.01718  -0.0930   0.9671   0.0070
  -8.750  -0.5195   0.02000   0.01557  -0.0944   0.9655   0.0073
  -8.500  -0.4972   0.01892   0.01432  -0.0940   0.9604   0.0077
  -8.250  -0.4696   0.01780   0.01300  -0.0947   0.9573   0.0081
  -8.000  -0.4403   0.01674   0.01174  -0.0956   0.9549   0.0087
  -7.750  -0.4096   0.01584   0.01061  -0.0967   0.9529   0.0093
  -7.500  -0.3862   0.01492   0.00955  -0.0962   0.9475   0.0102
  -7.250  -0.3577   0.01438   0.00893  -0.0966   0.9436   0.0109
  -7.000  -0.3270   0.01382   0.00828  -0.0975   0.9406   0.0119
  -6.750  -0.2964   0.01328   0.00762  -0.0982   0.9376   0.0131
  -6.500  -0.2712   0.01274   0.00699  -0.0978   0.9318   0.0144
  -6.250  -0.2406   0.01261   0.00688  -0.0985   0.9277   0.0163
  -6.000  -0.2089   0.01233   0.00651  -0.0994   0.9242   0.0183
  -5.750  -0.1829   0.01207   0.00617  -0.0990   0.9178   0.0198
  -5.500  -0.1546   0.01158   0.00560  -0.0993   0.9125   0.0216
  -5.250  -0.1261   0.01141   0.00541  -0.0995   0.9069   0.0238
  -5.000  -0.0988   0.01115   0.00509  -0.0994   0.9000   0.0257
  -4.750  -0.0703   0.01092   0.00477  -0.0995   0.8936   0.0277
  -4.500  -0.0436   0.01072   0.00450  -0.0993   0.8856   0.0289
  -4.250  -0.0159   0.01054   0.00424  -0.0992   0.8782   0.0296
  -4.000   0.0095   0.01005   0.00366  -0.0988   0.8700   0.0314
  -3.750   0.0356   0.00969   0.00324  -0.0984   0.8622   0.0332
  -3.500   0.0624   0.00944   0.00294  -0.0982   0.8540   0.0352
  -3.250   0.0887   0.00924   0.00268  -0.0978   0.8448   0.0366
  -3.000   0.1156   0.00907   0.00243  -0.0975   0.8357   0.0381
  -2.750   0.1424   0.00893   0.00221  -0.0973   0.8255   0.0395
  -2.500   0.1690   0.00881   0.00203  -0.0969   0.8143   0.0408
  -2.250   0.1956   0.00870   0.00185  -0.0966   0.8025   0.0436
  -2.000   0.2219   0.00858   0.00171  -0.0962   0.7896   0.0504
  -1.750   0.2477   0.00840   0.00159  -0.0958   0.7761   0.0825
  -1.500   0.2732   0.00821   0.00149  -0.0954   0.7618   0.1288
  -1.000   0.3227   0.00774   0.00138  -0.0943   0.7320   0.2887
  -0.750   0.3478   0.00763   0.00134  -0.0938   0.7170   0.3382
  -0.500   0.3729   0.00754   0.00132  -0.0933   0.7019   0.3834
  -0.250   0.3973   0.00739   0.00133  -0.0927   0.6867   0.4569
   0.000   0.4207   0.00720   0.00136  -0.0918   0.6712   0.5531
   0.250   0.4435   0.00704   0.00141  -0.0908   0.6553   0.6497
   0.500   0.4635   0.00682   0.00148  -0.0890   0.6360   0.7606
   0.750   0.4944   0.00660   0.00162  -0.0891   0.6126   0.9437
   1.000   0.5375   0.00680   0.00165  -0.0926   0.5816   0.9779
   1.250   0.5720   0.00698   0.00170  -0.0942   0.5562   0.9929
   1.500   0.6000   0.00719   0.00176  -0.0943   0.5326   1.0000
   1.750   0.6223   0.00736   0.00183  -0.0932   0.5122   1.0000
   2.000   0.6452   0.00753   0.00192  -0.0922   0.4953   1.0000
   2.250   0.6685   0.00769   0.00200  -0.0913   0.4800   1.0000
   2.500   0.6921   0.00785   0.00210  -0.0905   0.4654   1.0000
   2.750   0.7156   0.00803   0.00221  -0.0896   0.4499   1.0000
   3.000   0.7395   0.00820   0.00233  -0.0888   0.4360   1.0000
   3.250   0.7635   0.00838   0.00245  -0.0881   0.4225   1.0000
   3.500   0.7870   0.00858   0.00260  -0.0873   0.4060   1.0000
   3.750   0.8104   0.00881   0.00276  -0.0865   0.3887   1.0000
   4.000   0.8345   0.00900   0.00292  -0.0858   0.3753   1.0000
   4.250   0.8583   0.00921   0.00309  -0.0851   0.3601   1.0000
   4.500   0.8811   0.00949   0.00330  -0.0842   0.3364   1.0000
   4.750   0.9014   0.00995   0.00356  -0.0829   0.2959   1.0000
   5.000   0.9195   0.01058   0.00390  -0.0813   0.2420   1.0000
   5.250   0.9383   0.01120   0.00428  -0.0798   0.1952   1.0000
   5.500   0.9550   0.01199   0.00476  -0.0781   0.1362   1.0000
   5.750   0.9715   0.01279   0.00530  -0.0763   0.0905   1.0000
   6.000   0.9926   0.01322   0.00568  -0.0752   0.0775   1.0000
   6.250   1.0087   0.01404   0.00626  -0.0733   0.0308   1.0000
   6.500   1.0280   0.01459   0.00681  -0.0719   0.0195   1.0000
   6.750   1.0489   0.01501   0.00729  -0.0707   0.0161   1.0000
   7.000   1.0694   0.01542   0.00776  -0.0695   0.0142   1.0000
   7.250   1.0887   0.01592   0.00829  -0.0681   0.0123   1.0000
   7.500   1.1064   0.01652   0.00897  -0.0664   0.0107   1.0000
   7.750   1.1249   0.01701   0.00952  -0.0649   0.0101   1.0000
   8.000   1.1415   0.01753   0.01011  -0.0630   0.0094   1.0000
   8.250   1.1566   0.01812   0.01077  -0.0609   0.0088   1.0000
   8.500   1.1713   0.01871   0.01141  -0.0588   0.0083   1.0000
   8.750   1.1834   0.01946   0.01222  -0.0563   0.0077   1.0000
   9.000   1.1938   0.02033   0.01316  -0.0537   0.0072   1.0000
   9.250   1.2064   0.02106   0.01397  -0.0514   0.0068   1.0000
   9.500   1.2176   0.02188   0.01487  -0.0491   0.0065   1.0000
   9.750   1.2271   0.02282   0.01591  -0.0466   0.0062   1.0000
  10.000   1.2363   0.02381   0.01697  -0.0442   0.0060   1.0000
  10.250   1.2455   0.02482   0.01805  -0.0420   0.0058   1.0000
  10.500   1.2528   0.02598   0.01929  -0.0397   0.0056   1.0000
  10.750   1.2599   0.02720   0.02062  -0.0375   0.0054   1.0000
  11.000   1.2651   0.02861   0.02210  -0.0353   0.0052   1.0000
  11.250   1.2669   0.03036   0.02395  -0.0329   0.0051   1.0000
  11.500   1.2668   0.03238   0.02607  -0.0306   0.0050   1.0000
  11.750   1.2724   0.03400   0.02782  -0.0290   0.0048   1.0000
  12.000   1.2770   0.03578   0.02971  -0.0275   0.0047   1.0000
  12.250   1.2818   0.03760   0.03165  -0.0262   0.0045   1.0000
  12.500   1.2845   0.03970   0.03388  -0.0249   0.0043   1.0000
  12.750   1.2868   0.04193   0.03624  -0.0238   0.0042   1.0000
  13.000   1.2878   0.04438   0.03882  -0.0229   0.0041   1.0000
  13.250   1.2871   0.04715   0.04172  -0.0221   0.0040   1.0000
  13.500   1.2883   0.04974   0.04443  -0.0217   0.0039   1.0000
  13.750   1.2876   0.05266   0.04748  -0.0215   0.0038   1.0000
  14.000   1.2852   0.05593   0.05089  -0.0214   0.0038   1.0000
  14.250   1.2820   0.05942   0.05456  -0.0216   0.0038   1.0000
  14.500   1.2818   0.06259   0.05781  -0.0223   0.0036   1.0000
  14.750   1.2787   0.06629   0.06163  -0.0232   0.0035   1.0000
  15.000   1.2722   0.07066   0.06616  -0.0241   0.0035   1.0000
  15.250   1.2670   0.07499   0.07063  -0.0255   0.0035   1.0000
  15.500   1.2599   0.07977   0.07555  -0.0271   0.0035   1.0000
  15.750   1.2522   0.08480   0.08072  -0.0290   0.0034   1.0000
  16.000   1.2432   0.09023   0.08629  -0.0313   0.0034   1.0000
  16.250   1.2336   0.09596   0.09214  -0.0339   0.0034   1.0000
  16.500   1.2218   0.10230   0.09867  -0.0368   0.0034   1.0000
  16.750   1.2082   0.10916   0.10566  -0.0403   0.0033   1.0000
  17.000   1.1960   0.11602   0.11269  -0.0439   0.0033   1.0000
  17.250   1.1835   0.12321   0.12004  -0.0478   0.0033   1.0000
  17.500   1.1681   0.13120   0.12817  -0.0523   0.0033   1.0000
  17.750   1.1554   0.13896   0.13608  -0.0568   0.0033   1.0000
  18.000   1.1409   0.14743   0.14471  -0.0619   0.0033   1.0000
  18.250   1.1253   0.15657   0.15399  -0.0674   0.0033   1.0000
  18.500   1.1068   0.16701   0.16460  -0.0737   0.0034   1.0000
<< Back to GOE 602 AIRFOIL (goe602-il)

Polar data table (+)

Polar graphs


<< Back to GOE 602 AIRFOIL (goe602-il)