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

GOE 596 AIRFOIL (goe596-il) Xfoil prediction polar at RE=100,000 Ncrit=5


Details Polar file
Airfoil: GOE 596 AIRFOIL (goe596-il)
Reynolds number: 100,000
Max Cl/Cd: 57.45 at α=6.75°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-goe596-il-100000-n5.txt
Download as CSV file: xf-goe596-il-100000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 596 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
  ------ -------- --------- --------- -------- -------- --------
  -9.000  -0.3287   0.09965   0.09449  -0.0339   1.0000   0.0619
  -8.750  -0.3289   0.09700   0.09190  -0.0345   1.0000   0.0636
  -8.500  -0.3351   0.09464   0.08962  -0.0357   1.0000   0.0658
  -8.250  -0.2730   0.08245   0.07792  -0.0360   1.0000   0.0699
  -8.000  -0.2722   0.08002   0.07556  -0.0342   1.0000   0.0714
  -7.750  -0.2784   0.07787   0.07349  -0.0325   1.0000   0.0727
  -7.500  -0.3896   0.08516   0.08055  -0.0398   1.0000   0.0681
  -7.000  -0.3798   0.08069   0.07620  -0.0307   0.9999   0.0713
  -6.750  -0.3564   0.07640   0.07188  -0.0364   0.9923   0.0737
  -6.500  -0.3244   0.06965   0.06472  -0.0558   0.9772   0.0808
  -6.250  -0.3030   0.06412   0.05909  -0.0602   0.9695   0.0816
  -6.000  -0.2816   0.06058   0.05564  -0.0603   0.9638   0.0832
  -5.750  -0.2523   0.05031   0.04472  -0.0683   0.9535   0.0550
  -5.500  -0.2238   0.04652   0.04081  -0.0715   0.9477   0.0535
  -5.250  -0.1992   0.04268   0.03671  -0.0737   0.9381   0.0523
  -5.000  -0.1672   0.03859   0.03223  -0.0769   0.9320   0.0517
  -4.750  -0.1418   0.03556   0.02880  -0.0779   0.9215   0.0527
  -4.500  -0.1101   0.03256   0.02534  -0.0798   0.9146   0.0528
  -4.250  -0.0835   0.03017   0.02258  -0.0801   0.9043   0.0524
  -4.000  -0.0536   0.02802   0.02003  -0.0809   0.8957   0.0523
  -3.750  -0.0230   0.02616   0.01780  -0.0814   0.8865   0.0523
  -3.500   0.0056   0.02467   0.01596  -0.0815   0.8758   0.0529
  -3.250   0.0377   0.02348   0.01439  -0.0820   0.8667   0.0547
  -3.000   0.0672   0.02247   0.01308  -0.0820   0.8555   0.0555
  -2.750   0.0960   0.02141   0.01182  -0.0819   0.8436   0.0559
  -2.500   0.1253   0.02026   0.01056  -0.0820   0.8322   0.0566
  -2.250   0.1553   0.01934   0.00956  -0.0821   0.8210   0.0575
  -2.000   0.1827   0.01864   0.00880  -0.0817   0.8077   0.0586
  -1.750   0.2098   0.01805   0.00817  -0.0813   0.7941   0.0602
  -1.500   0.2368   0.01758   0.00766  -0.0809   0.7804   0.0633
  -1.250   0.2637   0.01716   0.00716  -0.0804   0.7669   0.0663
  -1.000   0.2902   0.01673   0.00667  -0.0798   0.7535   0.0684
  -0.750   0.3164   0.01632   0.00622  -0.0792   0.7400   0.0715
  -0.500   0.3429   0.01604   0.00585  -0.0786   0.7261   0.0759
  -0.250   0.3693   0.01580   0.00552  -0.0780   0.7118   0.0828
   0.000   0.3957   0.01551   0.00525  -0.0775   0.6979   0.1031
   0.250   0.4166   0.01387   0.00520  -0.0766   0.6856   0.5518
   0.750   0.5106   0.01317   0.00494  -0.0832   0.6621   1.0000
   1.000   0.5355   0.01333   0.00492  -0.0824   0.6517   1.0000
   1.250   0.5604   0.01351   0.00496  -0.0817   0.6414   1.0000
   1.500   0.5857   0.01369   0.00499  -0.0810   0.6324   1.0000
   1.750   0.6104   0.01388   0.00509  -0.0803   0.6224   1.0000
   2.000   0.6356   0.01409   0.00520  -0.0796   0.6135   1.0000
   2.250   0.6606   0.01430   0.00532  -0.0790   0.6045   1.0000
   2.500   0.6855   0.01453   0.00549  -0.0783   0.5955   1.0000
   2.750   0.7108   0.01475   0.00563  -0.0777   0.5871   1.0000
   3.000   0.7353   0.01499   0.00586  -0.0770   0.5777   1.0000
   3.250   0.7608   0.01523   0.00602  -0.0765   0.5698   1.0000
   3.500   0.7849   0.01548   0.00631  -0.0757   0.5601   1.0000
   3.750   0.8100   0.01574   0.00653  -0.0751   0.5518   1.0000
   4.000   0.8342   0.01599   0.00681  -0.0744   0.5422   1.0000
   4.250   0.8587   0.01626   0.00709  -0.0737   0.5330   1.0000
   4.500   0.8830   0.01650   0.00732  -0.0729   0.5228   1.0000
   4.750   0.9062   0.01675   0.00761  -0.0720   0.5107   1.0000
   5.000   0.9293   0.01699   0.00788  -0.0711   0.4983   1.0000
   5.250   0.9525   0.01723   0.00815  -0.0701   0.4861   1.0000
   5.500   0.9760   0.01749   0.00842  -0.0692   0.4751   1.0000
   5.750   0.9988   0.01778   0.00881  -0.0683   0.4634   1.0000
   6.000   1.0212   0.01804   0.00915  -0.0673   0.4500   1.0000
   6.250   1.0425   0.01829   0.00943  -0.0660   0.4338   1.0000
   6.500   1.0628   0.01854   0.00966  -0.0646   0.4147   1.0000
   6.750   1.0823   0.01884   0.00998  -0.0631   0.3938   1.0000
   7.000   1.1019   0.01921   0.01036  -0.0617   0.3746   1.0000
   7.250   1.1204   0.01964   0.01075  -0.0601   0.3545   1.0000
   7.500   1.1388   0.02012   0.01124  -0.0586   0.3356   1.0000
   7.750   1.1568   0.02064   0.01181  -0.0571   0.3165   1.0000
   8.000   1.1726   0.02126   0.01240  -0.0552   0.2931   1.0000
   8.250   1.1864   0.02196   0.01306  -0.0532   0.2624   1.0000
   8.500   1.1970   0.02285   0.01382  -0.0507   0.2210   1.0000
   8.750   1.2028   0.02406   0.01477  -0.0478   0.1790   1.0000
   9.000   1.2062   0.02545   0.01595  -0.0446   0.1419   1.0000
   9.250   1.2060   0.02698   0.01728  -0.0411   0.1033   1.0000
   9.500   1.2054   0.02860   0.01872  -0.0377   0.0847   1.0000
   9.750   1.2067   0.03017   0.02026  -0.0348   0.0759   1.0000
  10.250   1.2097   0.03348   0.02368  -0.0299   0.0672   1.0000
  10.500   1.2112   0.03527   0.02558  -0.0279   0.0641   1.0000
  10.750   1.2104   0.03735   0.02773  -0.0260   0.0614   1.0000
  11.250   1.2096   0.04179   0.03240  -0.0232   0.0571   1.0000
  11.500   1.2113   0.04401   0.03475  -0.0221   0.0551   1.0000
  11.750   1.2127   0.04633   0.03717  -0.0212   0.0530   1.0000
  12.000   1.2137   0.04879   0.03970  -0.0205   0.0511   1.0000
  12.250   1.2141   0.05135   0.04227  -0.0197   0.0493   1.0000
  12.500   1.2194   0.05352   0.04453  -0.0189   0.0476   1.0000
  12.750   1.2256   0.05566   0.04683  -0.0182   0.0459   1.0000
  13.000   1.2310   0.05793   0.04924  -0.0177   0.0439   1.0000
  13.250   1.2361   0.06025   0.05164  -0.0173   0.0421   1.0000
  13.500   1.2426   0.06246   0.05386  -0.0168   0.0405   1.0000
  13.750   1.2529   0.06450   0.05593  -0.0159   0.0390   1.0000
  14.000   1.2562   0.06726   0.05894  -0.0158   0.0377   1.0000
  14.250   1.2584   0.07019   0.06208  -0.0158   0.0363   1.0000
  14.500   1.2591   0.07332   0.06538  -0.0160   0.0350   1.0000
  14.750   1.2609   0.07638   0.06859  -0.0163   0.0339   1.0000
  15.000   1.2636   0.07940   0.07171  -0.0165   0.0330   1.0000
  15.250   1.2698   0.08212   0.07446  -0.0164   0.0320   1.0000
  15.500   1.2641   0.08641   0.07898  -0.0174   0.0313   1.0000
  15.750   1.2536   0.09145   0.08432  -0.0192   0.0308   1.0000
  16.000   1.2416   0.09690   0.09004  -0.0214   0.0302   1.0000
  16.250   1.2293   0.10264   0.09603  -0.0239   0.0298   1.0000
  16.500   1.2148   0.10899   0.10263  -0.0271   0.0293   1.0000
  16.750   1.1996   0.11578   0.10964  -0.0307   0.0291   1.0000
  17.000   1.1820   0.12335   0.11744  -0.0351   0.0288   1.0000
  17.250   1.1611   0.13210   0.12642  -0.0405   0.0289   1.0000
  17.500   1.1349   0.14280   0.13734  -0.0474   0.0291   1.0000
  17.750   1.1000   0.15691   0.15165  -0.0567   0.0296   1.0000
  18.000   1.0561   0.17606   0.17089  -0.0686   0.0305   1.0000
<< Back to GOE 596 AIRFOIL (goe596-il)

Polar data table (+)

Polar graphs


<< Back to GOE 596 AIRFOIL (goe596-il)