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

GOE 683 AIRFOIL (goe683-il) Xfoil prediction polar at RE=50,000 Ncrit=5


Details Polar file
Airfoil: GOE 683 AIRFOIL (goe683-il)
Reynolds number: 50,000
Max Cl/Cd: 21.07 at α=2.75°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-goe683-il-50000-n5.txt
Download as CSV file: xf-goe683-il-50000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 683 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
  ------ -------- --------- --------- -------- -------- --------
 -13.000  -0.3578   0.11318   0.10501  -0.0535   1.0000   0.1121
 -12.750  -0.3800   0.10582   0.09769  -0.0565   1.0000   0.1129
 -12.500  -0.3962   0.09976   0.09170  -0.0587   1.0000   0.1144
 -12.250  -0.3914   0.09776   0.08975  -0.0583   1.0000   0.1161
 -12.000  -0.4093   0.09173   0.08378  -0.0603   1.0000   0.1176
 -11.750  -0.5867   0.06578   0.05754  -0.0725   1.0000   0.1163
 -11.500  -0.6385   0.06107   0.05272  -0.0696   1.0000   0.1164
 -11.250  -0.6774   0.05849   0.05008  -0.0646   1.0000   0.1166
 -11.000  -0.7173   0.05697   0.04851  -0.0578   1.0000   0.1167
 -10.750  -0.7496   0.05578   0.04725  -0.0515   0.9982   0.1171
 -10.500  -0.7394   0.05326   0.04452  -0.0524   0.9832   0.1194
 -10.250  -0.7349   0.05073   0.04167  -0.0519   0.9681   0.1217
 -10.000  -0.7329   0.04821   0.03871  -0.0506   0.9532   0.1245
  -9.750  -0.7078   0.04682   0.03730  -0.0516   0.9418   0.1275
  -9.500  -0.6785   0.04528   0.03565  -0.0533   0.9327   0.1313
  -9.250  -0.6622   0.04355   0.03362  -0.0528   0.9201   0.1352
  -9.000  -0.6354   0.04191   0.03179  -0.0538   0.9109   0.1396
  -8.750  -0.6056   0.04076   0.03060  -0.0550   0.9010   0.1445
  -8.500  -0.5791   0.03921   0.02872  -0.0557   0.8915   0.1507
  -8.250  -0.5484   0.03818   0.02772  -0.0568   0.8817   0.1565
  -8.000  -0.5222   0.03710   0.02646  -0.0571   0.8717   0.1642
  -7.750  -0.4931   0.03611   0.02545  -0.0577   0.8617   0.1720
  -7.500  -0.4698   0.03523   0.02438  -0.0573   0.8511   0.1819
  -7.250  -0.4424   0.03447   0.02367  -0.0574   0.8412   0.1922
  -7.000  -0.4227   0.03384   0.02295  -0.0562   0.8301   0.2035
  -6.750  -0.3983   0.03317   0.02211  -0.0557   0.8206   0.2175
  -6.500  -0.3818   0.03284   0.02186  -0.0538   0.8092   0.2284
  -6.250  -0.3563   0.03228   0.02122  -0.0533   0.8002   0.2422
  -6.000  -0.3419   0.03203   0.02083  -0.0510   0.7889   0.2546
  -5.750  -0.3149   0.03156   0.02035  -0.0508   0.7805   0.2685
  -5.500  -0.2969   0.03133   0.02009  -0.0492   0.7699   0.2804
  -5.250  -0.2723   0.03094   0.01962  -0.0485   0.7613   0.2963
  -5.000  -0.2564   0.03079   0.01944  -0.0464   0.7513   0.3120
  -4.750  -0.2323   0.03046   0.01914  -0.0457   0.7427   0.3295
  -4.500  -0.2082   0.03017   0.01887  -0.0450   0.7340   0.3471
  -4.250  -0.1847   0.02989   0.01864  -0.0443   0.7246   0.3649
  -4.000  -0.1506   0.02938   0.01813  -0.0452   0.7180   0.3855
  -3.750  -0.1317   0.02931   0.01813  -0.0438   0.7074   0.4025
  -3.500  -0.0973   0.02881   0.01765  -0.0447   0.7007   0.4258
  -3.250  -0.0788   0.02874   0.01768  -0.0432   0.6909   0.4492
  -3.000  -0.0506   0.02839   0.01743  -0.0430   0.6834   0.4804
  -2.750  -0.0237   0.02811   0.01728  -0.0427   0.6760   0.5177
  -2.500   0.0019   0.02801   0.01742  -0.0421   0.6670   0.5609
  -2.250   0.0575   0.02757   0.01720  -0.0463   0.6609   0.6253
  -2.000   0.1109   0.02794   0.01778  -0.0504   0.6513   0.6860
  -1.750   0.1793   0.02845   0.01830  -0.0562   0.6436   0.7370
  -1.500   0.2294   0.02911   0.01887  -0.0590   0.6362   0.7740
  -1.250   0.2626   0.02987   0.01957  -0.0592   0.6273   0.8063
  -1.000   0.3039   0.03029   0.01980  -0.0606   0.6212   0.8354
  -0.750   0.3327   0.03110   0.02055  -0.0606   0.6126   0.8622
  -0.500   0.3705   0.03159   0.02091  -0.0621   0.6051   0.8858
  -0.250   0.4111   0.03181   0.02092  -0.0639   0.5999   0.9081
   0.000   0.4422   0.03261   0.02172  -0.0652   0.5903   0.9277
   0.250   0.4899   0.03275   0.02172  -0.0689   0.5837   0.9451
   0.500   0.5393   0.03290   0.02174  -0.0732   0.5771   0.9634
   0.750   0.5864   0.03311   0.02192  -0.0778   0.5685   0.9813
   1.000   0.6436   0.03257   0.02123  -0.0837   0.5627   0.9952
   1.250   0.6664   0.03281   0.02143  -0.0840   0.5560   1.0000
   1.500   0.6734   0.03332   0.02194  -0.0813   0.5492   1.0000
   1.750   0.6897   0.03345   0.02196  -0.0797   0.5443   1.0000
   2.000   0.7034   0.03376   0.02219  -0.0777   0.5393   1.0000
   2.250   0.7033   0.03475   0.02325  -0.0740   0.5320   1.0000
   2.500   0.7171   0.03509   0.02353  -0.0720   0.5269   1.0000
   2.750   0.7393   0.03509   0.02341  -0.0710   0.5231   1.0000
   3.000   0.7316   0.03653   0.02493  -0.0664   0.5157   1.0000
   3.250   0.7382   0.03729   0.02569  -0.0635   0.5101   1.0000
   3.500   0.7577   0.03747   0.02579  -0.0621   0.5061   1.0000
   3.750   0.7663   0.03820   0.02650  -0.0595   0.5012   1.0000
   4.000   0.7474   0.04020   0.02859  -0.0535   0.4937   1.0000
   4.250   0.7607   0.04070   0.02906  -0.0514   0.4893   1.0000
   4.500   0.7871   0.04064   0.02891  -0.0509   0.4861   1.0000
   4.750   0.7352   0.04421   0.03261  -0.0413   0.4773   1.0000
   5.000   0.7304   0.04549   0.03388  -0.0371   0.4718   1.0000
   5.250   0.7562   0.04550   0.03383  -0.0365   0.4687   1.0000
   5.500   0.7911   0.04517   0.03343  -0.0369   0.4663   1.0000
   6.000   0.6884   0.05279   0.04114  -0.0206   0.4494   1.0000
   7.500   0.6144   0.07147   0.05985  -0.0067   0.4092   1.0000
   8.000   0.5906   0.07850   0.06691  -0.0038   0.3959   1.0000
   8.250   0.6085   0.07930   0.06771  -0.0029   0.3937   1.0000
   8.750   0.5786   0.08732   0.07578  -0.0009   0.3804   1.0000
   9.000   0.5939   0.08841   0.07687   0.0000   0.3774   1.0000
   9.250   0.6132   0.08914   0.07761   0.0009   0.3752   1.0000
   9.500   0.5814   0.09517   0.08369   0.0013   0.3680   1.0000
   9.750   0.5840   0.09758   0.08613   0.0020   0.3632   1.0000
  10.000   0.5981   0.09884   0.08742   0.0027   0.3599   1.0000
  10.250   0.6188   0.09948   0.08808   0.0036   0.3575   1.0000
  10.500   0.5928   0.10494   0.09359   0.0036   0.3503   1.0000
  10.750   0.5998   0.10681   0.09549   0.0043   0.3449   1.0000
<< Back to GOE 683 AIRFOIL (goe683-il)

Polar data table (+)

Polar graphs


<< Back to GOE 683 AIRFOIL (goe683-il)