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GOE 559 AIRFOIL (goe559-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: GOE 559 AIRFOIL (goe559-il)
Reynolds number: 50,000
Max Cl/Cd: 30.09 at α=4°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe559-il-50000.txt
Download as CSV file: xf-goe559-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 559 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 =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.250  -0.3615   0.11249   0.10646  -0.0307   1.0000   0.1845
  -9.000  -0.3592   0.10948   0.10353  -0.0305   1.0000   0.1988
  -8.750  -0.3737   0.10855   0.10274  -0.0321   1.0000   0.2045
  -8.500  -0.3583   0.10414   0.09836  -0.0297   1.0000   0.2306
  -8.250  -0.3629   0.10204   0.09638  -0.0292   1.0000   0.2466
  -8.000  -0.3573   0.09894   0.09337  -0.0271   1.0000   0.2747
  -7.750  -0.3504   0.09613   0.09062  -0.0248   1.0000   0.3029
  -7.500  -0.3552   0.09414   0.08876  -0.0223   1.0000   0.3331
  -7.250  -0.3295   0.08999   0.08462  -0.0181   1.0000   0.3876
  -7.000  -0.3213   0.08738   0.08209  -0.0147   1.0000   0.4319
  -6.750  -0.3184   0.08516   0.07993  -0.0107   1.0000   0.4768
  -6.500  -0.2984   0.08183   0.07665  -0.0071   1.0000   0.5343
  -6.250  -0.2561   0.07633   0.07110  -0.0046   1.0000   0.6188
  -5.750  -0.2008   0.06792   0.06276  -0.0038   1.0000   0.7251
  -5.500  -0.1833   0.06443   0.05931  -0.0040   1.0000   0.7550
  -4.250  -0.2343   0.05863   0.05442   0.0119   1.0000   0.7822
  -4.000  -0.2680   0.06009   0.05611   0.0184   1.0000   0.7878
  -3.750  -0.3124   0.05905   0.05522   0.0090   0.9626   0.6688
  -3.500  -0.2801   0.04601   0.04085  -0.0507   0.9222   0.3450
  -3.250  -0.2095   0.04007   0.03317  -0.0630   0.9092   0.2271
  -3.000  -0.1547   0.03755   0.02941  -0.0663   0.8965   0.1736
  -2.750  -0.1039   0.03545   0.02652  -0.0684   0.8847   0.1406
  -2.500  -0.0509   0.03384   0.02412  -0.0706   0.8739   0.1163
  -2.250   0.0053   0.03200   0.02185  -0.0734   0.8648   0.0992
  -2.000   0.0467   0.03099   0.02050  -0.0742   0.8531   0.0890
  -1.750   0.0852   0.02996   0.01926  -0.0751   0.8418   0.0847
  -1.500   0.1165   0.02919   0.01817  -0.0751   0.8308   0.0851
  -1.250   0.1525   0.02848   0.01710  -0.0760   0.8217   0.0880
  -1.000   0.1827   0.02816   0.01649  -0.0762   0.8115   0.0927
  -0.750   0.2060   0.02792   0.01621  -0.0756   0.8009   0.1160
  -0.500   0.3255   0.02452   0.01459  -0.0901   0.7962   1.0000
  -0.250   0.3401   0.02528   0.01496  -0.0884   0.7853   1.0000
   0.000   0.3581   0.02599   0.01531  -0.0874   0.7762   1.0000
   0.250   0.3852   0.02639   0.01537  -0.0875   0.7687   1.0000
   0.500   0.3944   0.02750   0.01627  -0.0855   0.7594   1.0000
   0.750   0.4260   0.02779   0.01630  -0.0862   0.7533   1.0000
   1.000   0.4285   0.02926   0.01770  -0.0835   0.7443   1.0000
   1.250   0.4570   0.02977   0.01809  -0.0840   0.7391   1.0000
   1.500   0.4529   0.03165   0.01994  -0.0808   0.7312   1.0000
   1.750   0.4815   0.03223   0.02050  -0.0813   0.7262   1.0000
   2.000   0.4741   0.03437   0.02266  -0.0779   0.7190   1.0000
   2.250   0.4896   0.03561   0.02393  -0.0772   0.7141   1.0000
   2.500   0.5055   0.03695   0.02532  -0.0766   0.7101   1.0000
   2.750   0.4864   0.03976   0.02809  -0.0725   0.7054   1.0000
   3.000   0.4886   0.04162   0.03001  -0.0705   0.7015   1.0000
   3.250   0.5089   0.04291   0.03144  -0.0706   0.6982   1.0000
   3.500   0.5140   0.04491   0.03359  -0.0692   0.6966   1.0000
   3.750   0.5052   0.04736   0.03608  -0.0666   0.6958   1.0000
   4.000   0.6836   0.02272   0.01003  -0.0496   0.2808   1.0000
   4.250   0.6813   0.02550   0.01175  -0.0460   0.0780   1.0000
   4.500   0.6964   0.02671   0.01316  -0.0439   0.0703   1.0000
   4.750   0.7111   0.02780   0.01454  -0.0419   0.0695   1.0000
   5.000   0.7239   0.02900   0.01603  -0.0396   0.0699   1.0000
   5.250   0.7346   0.03026   0.01761  -0.0372   0.0715   1.0000
   5.500   0.7416   0.03172   0.01934  -0.0344   0.0735   1.0000
   5.750   0.7440   0.03332   0.02118  -0.0312   0.0755   1.0000
   6.000   0.7429   0.03506   0.02307  -0.0280   0.0769   1.0000
   6.250   0.7382   0.03719   0.02535  -0.0250   0.0784   1.0000
   6.500   0.7461   0.03860   0.02703  -0.0230   0.0810   1.0000
   6.750   0.7529   0.04019   0.02882  -0.0208   0.0851   1.0000
   7.000   0.7594   0.04185   0.03056  -0.0184   0.0888   1.0000
   7.250   0.7812   0.04244   0.03127  -0.0156   0.0939   1.0000
   7.500   0.9674   0.03975   0.02899  -0.0208   0.1332   1.0000
   7.750   1.0720   0.04061   0.03059  -0.0247   0.1731   1.0000
   8.000   1.1217   0.04221   0.03322  -0.0242   0.2026   1.0000
   8.750   1.0784   0.03947   0.03324  -0.0118   0.2420   1.0000
   9.000   1.0503   0.04426   0.03840  -0.0080   0.2432   1.0000
   9.250   1.0209   0.04903   0.04339  -0.0047   0.2432   1.0000
   9.500   0.9915   0.05372   0.04817  -0.0018   0.2424   1.0000
   9.750   0.9635   0.05904   0.05355  -0.0003   0.2414   1.0000
  10.000   0.9157   0.06351   0.05802   0.0010   0.2296   1.0000
  10.250   0.8719   0.06973   0.06422   0.0003   0.2196   1.0000
  10.500   0.8424   0.07662   0.07110  -0.0012   0.2158   1.0000
  10.750   0.8001   0.08425   0.07867  -0.0038   0.2063   1.0000
  11.000   0.7692   0.09216   0.08654  -0.0068   0.2028   1.0000
  11.250   0.7415   0.10004   0.09437  -0.0100   0.2016   1.0000
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