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

GOE 398 AIRFOIL (goe398-il) Xfoil prediction polar at RE=50,000 Ncrit=5


Details Polar file
Airfoil: GOE 398 AIRFOIL (goe398-il)
Reynolds number: 50,000
Max Cl/Cd: 31.01 at α=7°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-goe398-il-50000-n5.txt
Download as CSV file: xf-goe398-il-50000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 398 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
  ------ -------- --------- --------- -------- -------- --------
  -9.500  -0.3150   0.10712   0.10032  -0.0401   1.0000   0.0825
  -9.250  -0.3281   0.10483   0.09814  -0.0386   1.0000   0.0809
  -8.750  -0.3917   0.09843   0.09201  -0.0379   1.0000   0.0760
  -8.500  -0.3883   0.09360   0.08719  -0.0429   0.9924   0.0760
  -8.250  -0.3867   0.08796   0.08153  -0.0494   0.9827   0.0760
  -8.000  -0.3833   0.08132   0.07481  -0.0567   0.9725   0.0760
  -7.750  -0.3810   0.07513   0.06849  -0.0621   0.9619   0.0760
  -7.500  -0.3778   0.06967   0.06286  -0.0658   0.9520   0.0758
  -7.250  -0.3712   0.06375   0.05663  -0.0701   0.9437   0.0758
  -7.000  -0.3660   0.05953   0.05218  -0.0714   0.9338   0.0767
  -6.750  -0.3418   0.05751   0.05008  -0.0730   0.9282   0.0788
  -6.500  -0.3334   0.05447   0.04680  -0.0730   0.9187   0.0802
  -6.250  -0.3138   0.05044   0.04233  -0.0752   0.9128   0.0816
  -6.000  -0.3047   0.04748   0.03897  -0.0743   0.9034   0.0825
  -5.750  -0.2827   0.04428   0.03519  -0.0753   0.8974   0.0849
  -5.500  -0.2664   0.04180   0.03211  -0.0746   0.8898   0.0874
  -5.250  -0.2427   0.04040   0.03062  -0.0746   0.8830   0.0897
  -5.000  -0.2108   0.03868   0.02861  -0.0758   0.8783   0.0925
  -4.750  -0.1942   0.03755   0.02719  -0.0743   0.8696   0.0958
  -4.500  -0.1641   0.03602   0.02524  -0.0748   0.8634   0.1003
  -4.250  -0.1336   0.03498   0.02415  -0.0753   0.8570   0.1045
  -4.000  -0.1088   0.03405   0.02300  -0.0747   0.8478   0.1100
  -3.750  -0.0687   0.03290   0.02173  -0.0766   0.8425   0.1180
  -3.500  -0.0487   0.03232   0.02101  -0.0751   0.8319   0.1255
  -3.250  -0.0103   0.03144   0.02011  -0.0766   0.8260   0.1371
  -3.000   0.0145   0.03099   0.01961  -0.0760   0.8171   0.1498
  -2.750   0.0485   0.03040   0.01898  -0.0768   0.8100   0.1690
  -2.250   0.1068   0.02942   0.01801  -0.0770   0.7944   0.2238
  -2.000   0.1453   0.02858   0.01735  -0.0786   0.7891   0.2706
  -1.750   0.1638   0.02830   0.01729  -0.0771   0.7789   0.3161
  -1.500   0.1959   0.02758   0.01697  -0.0775   0.7725   0.4025
  -1.250   0.2159   0.02705   0.01699  -0.0757   0.7635   0.5233
  -1.000   0.3412   0.02559   0.01647  -0.0923   0.7608   1.0000
  -0.750   0.3739   0.02549   0.01604  -0.0928   0.7536   1.0000
  -0.500   0.3891   0.02580   0.01614  -0.0905   0.7420   1.0000
  -0.250   0.4241   0.02563   0.01570  -0.0912   0.7351   1.0000
   0.000   0.4378   0.02600   0.01591  -0.0887   0.7229   1.0000
   0.250   0.4746   0.02579   0.01547  -0.0897   0.7166   1.0000
   0.500   0.4873   0.02620   0.01576  -0.0870   0.7037   1.0000
   0.750   0.5138   0.02627   0.01567  -0.0864   0.6947   1.0000
   1.000   0.5379   0.02640   0.01565  -0.0854   0.6846   1.0000
   1.250   0.5577   0.02667   0.01581  -0.0838   0.6735   1.0000
   1.500   0.5898   0.02656   0.01556  -0.0840   0.6652   1.0000
   1.750   0.6054   0.02699   0.01590  -0.0818   0.6531   1.0000
   2.000   0.6420   0.02675   0.01551  -0.0826   0.6460   1.0000
   2.250   0.6546   0.02729   0.01600  -0.0800   0.6330   1.0000
   2.500   0.6783   0.02749   0.01611  -0.0790   0.6230   1.0000
   2.750   0.7043   0.02760   0.01613  -0.0783   0.6132   1.0000
   3.000   0.7212   0.02805   0.01653  -0.0764   0.6016   1.0000
   3.250   0.7547   0.02790   0.01627  -0.0767   0.5935   1.0000
   3.500   0.7666   0.02852   0.01688  -0.0742   0.5810   1.0000
   3.750   0.7911   0.02871   0.01700  -0.0733   0.5713   1.0000
   4.000   0.8138   0.02893   0.01717  -0.0721   0.5608   1.0000
   4.250   0.8291   0.02946   0.01768  -0.0701   0.5496   1.0000
   4.500   0.8625   0.02927   0.01741  -0.0703   0.5411   1.0000
   4.750   0.8721   0.03003   0.01818  -0.0676   0.5292   1.0000
   5.000   0.8929   0.03036   0.01847  -0.0663   0.5192   1.0000
   5.250   0.9172   0.03056   0.01863  -0.0654   0.5096   1.0000
   5.500   0.9309   0.03121   0.01930  -0.0633   0.4990   1.0000
   5.750   0.9639   0.03110   0.01910  -0.0635   0.4907   1.0000
   6.000   0.9718   0.03200   0.02006  -0.0607   0.4797   1.0000
   6.250   0.9966   0.03226   0.02029  -0.0600   0.4707   1.0000
   6.500   1.0129   0.03287   0.02092  -0.0583   0.4610   1.0000
   6.750   1.0318   0.03343   0.02148  -0.0569   0.4519   1.0000
   7.000   1.0519   0.03392   0.02199  -0.0557   0.4427   1.0000
   7.250   1.0657   0.03471   0.02280  -0.0537   0.4335   1.0000
   7.500   1.0872   0.03522   0.02331  -0.0528   0.4248   1.0000
   7.750   1.0983   0.03616   0.02431  -0.0506   0.4161   1.0000
   8.000   1.1169   0.03683   0.02502  -0.0493   0.4076   1.0000
   8.250   1.1298   0.03780   0.02604  -0.0475   0.3996   1.0000
   8.500   1.1402   0.03887   0.02719  -0.0455   0.3915   1.0000
   8.750   1.1628   0.03948   0.02783  -0.0448   0.3844   1.0000
   9.000   1.1581   0.04129   0.02978  -0.0414   0.3760   1.0000
   9.250   1.1912   0.04139   0.02987  -0.0417   0.3696   1.0000
   9.500   1.1732   0.04404   0.03271  -0.0376   0.3617   1.0000
   9.750   1.1889   0.04504   0.03377  -0.0364   0.3553   1.0000
  10.000   1.2022   0.04626   0.03508  -0.0352   0.3496   1.0000
  10.250   1.1785   0.04983   0.03884  -0.0318   0.3422   1.0000
  10.500   1.2047   0.05015   0.03921  -0.0313   0.3371   1.0000
  10.750   1.1757   0.05464   0.04386  -0.0286   0.3303   1.0000
  11.000   1.1569   0.05876   0.04811  -0.0269   0.3233   1.0000
  11.250   1.1940   0.05795   0.04735  -0.0265   0.3197   1.0000
  11.500   1.0788   0.07342   0.06298  -0.0264   0.3070   1.0000
  11.750   1.1032   0.07350   0.06315  -0.0255   0.3041   1.0000
  12.250   1.0364   0.08984   0.07963  -0.0284   0.2877   1.0000
  12.750   0.9849   0.10590   0.09581  -0.0328   0.2728   1.0000
<< Back to GOE 398 AIRFOIL (goe398-il)

Polar data table (+)

Polar graphs


<< Back to GOE 398 AIRFOIL (goe398-il)