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

GOE 7K AIRFOIL (goe07k-il) Xfoil prediction polar at RE=50,000 Ncrit=5


Details Polar file
Airfoil: GOE 7K AIRFOIL (goe07k-il)
Reynolds number: 50,000
Max Cl/Cd: 25.18 at α=8.75°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-goe07k-il-50000-n5.txt
Download as CSV file: xf-goe07k-il-50000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 7K 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
  ------ -------- --------- --------- -------- -------- --------
  -8.250  -0.4729   0.12570   0.11961  -0.0266   1.0000   0.0863
  -8.000  -0.4856   0.12382   0.11781  -0.0286   0.9981   0.0875
  -7.750  -0.4939   0.12149   0.11554  -0.0320   0.9939   0.0879
  -7.500  -0.5013   0.11876   0.11283  -0.0362   0.9893   0.0883
  -7.250  -0.4714   0.11242   0.10645  -0.0308   0.9895   0.0938
  -7.000  -0.4736   0.10953   0.10358  -0.0317   0.9860   0.0961
  -6.750  -0.4760   0.10622   0.10029  -0.0339   0.9817   0.0992
  -6.500  -0.4865   0.10342   0.09741  -0.0430   0.9749   0.1035
  -6.250  -0.4778   0.09860   0.09266  -0.0410   0.9728   0.1052
  -6.000  -0.4689   0.09478   0.08883  -0.0406   0.9702   0.1074
  -5.750  -0.4563   0.08818   0.08202  -0.0447   0.9664   0.0688
  -5.250  -0.4305   0.07597   0.06922  -0.0517   0.9586   0.0472
  -5.000  -0.4207   0.07210   0.06518  -0.0518   0.9555   0.0469
  -4.750  -0.4098   0.06834   0.06116  -0.0517   0.9524   0.0470
  -4.500  -0.3947   0.06434   0.05676  -0.0521   0.9493   0.0473
  -4.250  -0.3772   0.06071   0.05280  -0.0523   0.9466   0.0471
  -4.000  -0.3592   0.05727   0.04897  -0.0522   0.9441   0.0469
  -3.750  -0.3454   0.05427   0.04555  -0.0507   0.9410   0.0468
  -3.500  -0.3286   0.05132   0.04201  -0.0493   0.9384   0.0482
  -3.250  -0.3111   0.04931   0.03992  -0.0487   0.9359   0.0512
  -3.000  -0.2886   0.04709   0.03720  -0.0480   0.9334   0.0531
  -2.750  -0.2631   0.04497   0.03447  -0.0475   0.9313   0.0562
  -2.500  -0.2432   0.04341   0.03224  -0.0457   0.9285   0.0596
  -2.250  -0.2234   0.04170   0.03026  -0.0443   0.9260   0.0630
  -2.000  -0.2006   0.04062   0.02888  -0.0435   0.9232   0.0691
  -1.750  -0.1726   0.03938   0.02727  -0.0435   0.9208   0.0765
  -1.500  -0.1427   0.03862   0.02618  -0.0438   0.9187   0.0873
  -1.250  -0.1070   0.03798   0.02529  -0.0456   0.9171   0.1025
  -1.000  -0.0786   0.03748   0.02464  -0.0461   0.9147   0.1158
  -0.750  -0.0547   0.03710   0.02412  -0.0458   0.9114   0.1323
  -0.500  -0.0277   0.03685   0.02381  -0.0460   0.9084   0.1491
  -0.250   0.0717   0.03483   0.02445  -0.0622   0.9112   1.0000
   0.000   0.0945   0.03524   0.02444  -0.0616   0.9075   1.0000
   0.250   0.1214   0.03573   0.02456  -0.0619   0.9041   1.0000
   0.500   0.1495   0.03627   0.02479  -0.0625   0.9009   1.0000
   0.750   0.1654   0.03658   0.02489  -0.0609   0.8957   1.0000
   1.000   0.1902   0.03705   0.02513  -0.0609   0.8913   1.0000
   1.250   0.2206   0.03763   0.02550  -0.0620   0.8878   1.0000
   1.500   0.2377   0.03798   0.02570  -0.0606   0.8816   1.0000
   1.750   0.2635   0.03845   0.02603  -0.0608   0.8764   1.0000
   2.000   0.2957   0.03904   0.02649  -0.0622   0.8728   1.0000
   2.250   0.3101   0.03937   0.02675  -0.0604   0.8652   1.0000
   2.500   0.3391   0.03989   0.02719  -0.0612   0.8604   1.0000
   2.750   0.3602   0.04035   0.02761  -0.0606   0.8541   1.0000
   3.000   0.3847   0.04082   0.02804  -0.0606   0.8479   1.0000
   3.250   0.4147   0.04132   0.02853  -0.0616   0.8429   1.0000
   3.500   0.4318   0.04175   0.02897  -0.0603   0.8347   1.0000
   3.750   0.4650   0.04223   0.02948  -0.0617   0.8300   1.0000
   4.000   0.4794   0.04269   0.02998  -0.0600   0.8210   1.0000
   4.250   0.5119   0.04313   0.03048  -0.0613   0.8159   1.0000
   4.750   0.5605   0.04399   0.03154  -0.0611   0.8012   1.0000
   5.000   0.5750   0.04450   0.03214  -0.0595   0.7909   1.0000
   5.250   0.6060   0.04488   0.03267  -0.0605   0.7851   1.0000
   5.500   0.6233   0.04538   0.03332  -0.0593   0.7749   1.0000
   5.750   0.6440   0.04586   0.03394  -0.0586   0.7657   1.0000
   6.000   0.6738   0.04613   0.03445  -0.0592   0.7583   1.0000
   6.250   0.6904   0.04666   0.03517  -0.0579   0.7471   1.0000
   6.500   0.7152   0.04700   0.03572  -0.0577   0.7379   1.0000
   6.750   0.7458   0.04692   0.03593  -0.0580   0.7281   1.0000
   7.000   0.7716   0.04658   0.03586  -0.0572   0.7136   1.0000
   7.250   0.7979   0.04609   0.03569  -0.0563   0.6986   1.0000
   7.500   0.8222   0.04574   0.03566  -0.0551   0.6835   1.0000
   7.750   0.8509   0.04482   0.03510  -0.0540   0.6672   1.0000
   8.000   0.8644   0.04427   0.03487  -0.0508   0.6413   1.0000
   8.250   0.8881   0.04216   0.03312  -0.0472   0.6076   1.0000
   8.500   0.8998   0.04145   0.03270  -0.0433   0.5663   1.0000
   8.750   0.9382   0.03726   0.02692  -0.0373   0.2935   1.0000
   9.000   0.9255   0.04014   0.02885  -0.0326   0.1838   1.0000
   9.250   0.9193   0.04299   0.03103  -0.0292   0.1139   1.0000
   9.500   0.9184   0.04554   0.03325  -0.0264   0.0859   1.0000
   9.750   0.9224   0.04776   0.03541  -0.0239   0.0711   1.0000
  10.000   0.9322   0.04958   0.03734  -0.0218   0.0607   1.0000
  10.250   0.9487   0.05110   0.03904  -0.0201   0.0533   1.0000
  10.500   0.9764   0.05242   0.04036  -0.0194   0.0464   1.0000
  10.750   1.0254   0.05360   0.04199  -0.0202   0.0403   1.0000
  11.000   1.0695   0.05600   0.04452  -0.0216   0.0363   1.0000
  11.250   1.1013   0.05921   0.04834  -0.0215   0.0342   1.0000
  11.500   1.1176   0.06279   0.05244  -0.0202   0.0331   1.0000
  11.750   1.1209   0.06615   0.05622  -0.0177   0.0322   1.0000
  12.000   1.1171   0.06949   0.05997  -0.0149   0.0315   1.0000
  12.250   1.1096   0.07281   0.06363  -0.0121   0.0310   1.0000
  12.500   1.0983   0.07638   0.06751  -0.0093   0.0307   1.0000
  12.750   1.0841   0.08013   0.07155  -0.0068   0.0304   1.0000
  13.000   1.0677   0.08418   0.07589  -0.0047   0.0305   1.0000
  13.250   1.0487   0.08849   0.08045  -0.0031   0.0303   1.0000
  13.500   1.0285   0.09314   0.08533  -0.0023   0.0304   1.0000
  13.750   1.0067   0.09833   0.09074  -0.0023   0.0307   1.0000
  14.000   0.9839   0.10411   0.09672  -0.0032   0.0311   1.0000
  14.250   0.9602   0.11070   0.10346  -0.0053   0.0315   1.0000
<< Back to GOE 7K AIRFOIL (goe07k-il)

Polar data table (+)

Polar graphs


<< Back to GOE 7K AIRFOIL (goe07k-il)