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

GOE 371 AIRFOIL (goe371-il) Xfoil prediction polar at RE=500,000 Ncrit=5


Details Polar file
Airfoil: GOE 371 AIRFOIL (goe371-il)
Reynolds number: 500,000
Max Cl/Cd: 103.93 at α=4.5°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-goe371-il-500000-n5.txt
Download as CSV file: xf-goe371-il-500000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 371 AIRFOIL                                 
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.500 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.750  -0.3218   0.09734   0.09509  -0.0283   1.0000   0.0068
  -8.500  -0.3211   0.09438   0.09217  -0.0285   1.0000   0.0069
  -8.250  -0.3220   0.09157   0.08939  -0.0285   1.0000   0.0071
  -8.000  -0.3187   0.08706   0.08491  -0.0314   0.9979   0.0076
  -7.500  -0.2905   0.07863   0.07649  -0.0409   0.9799   0.0076
  -7.250  -0.2696   0.07435   0.07221  -0.0474   0.9713   0.0079
  -7.000  -0.2433   0.06871   0.06655  -0.0565   0.9635   0.0079
  -6.750  -0.2146   0.06401   0.06180  -0.0647   0.9534   0.0081
  -6.500  -0.1849   0.05969   0.05743  -0.0723   0.9409   0.0085
  -6.250  -0.1556   0.05447   0.05212  -0.0804   0.9251   0.0088
  -6.000  -0.1294   0.04915   0.04667  -0.0873   0.9062   0.0094
  -5.750  -0.1038   0.03954   0.03677  -0.0963   0.8858   0.0107
  -5.500  -0.0797   0.03657   0.03361  -0.0986   0.8702   0.0112
  -5.250  -0.0553   0.03363   0.03048  -0.1004   0.8558   0.0117
  -5.000  -0.0346   0.01952   0.01525  -0.1038   0.8426   0.0150
  -4.750  -0.0085   0.01995   0.01564  -0.1036   0.8303   0.0155
  -4.500   0.0171   0.01960   0.01515  -0.1034   0.8193   0.0162
  -4.000   0.0678   0.01577   0.01042  -0.1026   0.8007   0.0204
  -3.750   0.0926   0.01445   0.00886  -0.1023   0.7924   0.0214
  -3.500   0.1186   0.01404   0.00833  -0.1020   0.7831   0.0222
  -3.250   0.1450   0.01376   0.00796  -0.1018   0.7744   0.0234
  -3.000   0.1713   0.01319   0.00721  -0.1014   0.7662   0.0246
  -2.750   0.1978   0.01251   0.00634  -0.1011   0.7585   0.0254
  -2.500   0.2245   0.01205   0.00572  -0.1007   0.7519   0.0265
  -2.250   0.2515   0.01169   0.00525  -0.1004   0.7454   0.0274
  -2.000   0.2783   0.01132   0.00476  -0.1001   0.7381   0.0279
  -1.750   0.3052   0.01104   0.00439  -0.0998   0.7311   0.0282
  -1.500   0.3320   0.01070   0.00397  -0.0995   0.7239   0.0286
  -1.250   0.3584   0.01014   0.00333  -0.0991   0.7172   0.0290
  -1.000   0.3847   0.00970   0.00284  -0.0987   0.7096   0.0296
  -0.750   0.4111   0.00939   0.00249  -0.0984   0.7019   0.0305
  -0.500   0.4376   0.00919   0.00225  -0.0980   0.6928   0.0306
  -0.250   0.4642   0.00901   0.00204  -0.0977   0.6823   0.0307
   0.000   0.4906   0.00887   0.00186  -0.0973   0.6695   0.0309
   0.250   0.5168   0.00877   0.00170  -0.0968   0.6518   0.0312
   0.500   0.5423   0.00873   0.00156  -0.0962   0.6238   0.0319
   0.750   0.5660   0.00885   0.00146  -0.0953   0.5761   0.0328
   1.000   0.5895   0.00905   0.00144  -0.0944   0.5349   0.0340
   1.250   0.6145   0.00919   0.00145  -0.0938   0.5112   0.0354
   1.500   0.6402   0.00930   0.00148  -0.0934   0.4947   0.0370
   1.750   0.6660   0.00941   0.00152  -0.0929   0.4798   0.0384
   2.000   0.6918   0.00953   0.00158  -0.0925   0.4655   0.0414
   2.250   0.7175   0.00952   0.00169  -0.0921   0.4535   0.1065
   2.750   0.7803   0.00803   0.00208  -0.0944   0.4334   1.0000
   3.000   0.8062   0.00818   0.00219  -0.0940   0.4254   1.0000
   3.250   0.8320   0.00834   0.00230  -0.0936   0.4158   1.0000
   3.500   0.8579   0.00848   0.00242  -0.0932   0.4072   1.0000
   3.750   0.8835   0.00865   0.00256  -0.0927   0.3979   1.0000
   4.000   0.9089   0.00884   0.00271  -0.0923   0.3840   1.0000
   4.250   0.9341   0.00903   0.00286  -0.0918   0.3687   1.0000
   4.500   0.9593   0.00923   0.00302  -0.0913   0.3526   1.0000
   4.750   0.9836   0.00950   0.00322  -0.0907   0.3264   1.0000
   5.000   1.0055   0.01000   0.00348  -0.0897   0.2774   1.0000
   5.250   1.0268   0.01057   0.00384  -0.0887   0.2353   1.0000
   5.500   1.0489   0.01107   0.00420  -0.0878   0.2072   1.0000
   5.750   1.0712   0.01154   0.00457  -0.0870   0.1841   1.0000
   6.000   1.0942   0.01195   0.00490  -0.0862   0.1675   1.0000
   6.250   1.1173   0.01233   0.00524  -0.0855   0.1523   1.0000
   6.500   1.1395   0.01279   0.00560  -0.0847   0.1290   1.0000
   6.750   1.1607   0.01334   0.00603  -0.0837   0.1011   1.0000
   7.000   1.1790   0.01417   0.00662  -0.0823   0.0652   1.0000
   7.250   1.1987   0.01483   0.00719  -0.0811   0.0461   1.0000
   7.500   1.2154   0.01578   0.00795  -0.0795   0.0185   1.0000
   7.750   1.2361   0.01632   0.00854  -0.0783   0.0135   1.0000
   8.000   1.2559   0.01692   0.00922  -0.0771   0.0114   1.0000
   8.250   1.2743   0.01765   0.01003  -0.0756   0.0095   1.0000
   8.500   1.2940   0.01820   0.01066  -0.0744   0.0089   1.0000
   8.750   1.3126   0.01882   0.01137  -0.0730   0.0082   1.0000
   9.000   1.3301   0.01951   0.01214  -0.0714   0.0076   1.0000
   9.250   1.3460   0.02028   0.01298  -0.0697   0.0070   1.0000
   9.500   1.3574   0.02135   0.01418  -0.0673   0.0064   1.0000
   9.750   1.3711   0.02205   0.01496  -0.0652   0.0061   1.0000
  10.000   1.3829   0.02282   0.01582  -0.0628   0.0057   1.0000
  10.250   1.3922   0.02374   0.01684  -0.0602   0.0055   1.0000
  10.500   1.4005   0.02476   0.01796  -0.0576   0.0052   1.0000
  10.750   1.4088   0.02580   0.01909  -0.0551   0.0050   1.0000
  11.000   1.4147   0.02705   0.02044  -0.0526   0.0049   1.0000
  11.250   1.4204   0.02835   0.02183  -0.0503   0.0047   1.0000
  11.500   1.4242   0.02987   0.02344  -0.0480   0.0045   1.0000
  11.750   1.4214   0.03202   0.02573  -0.0454   0.0043   1.0000
  12.000   1.4213   0.03409   0.02792  -0.0434   0.0043   1.0000
  12.250   1.4208   0.03632   0.03029  -0.0417   0.0042   1.0000
  12.500   1.4208   0.03864   0.03275  -0.0403   0.0041   1.0000
  12.750   1.4253   0.04061   0.03486  -0.0393   0.0039   1.0000
  13.000   1.4223   0.04346   0.03785  -0.0383   0.0039   1.0000
  13.250   1.4226   0.04608   0.04061  -0.0377   0.0037   1.0000
  13.500   1.4208   0.04903   0.04371  -0.0373   0.0036   1.0000
  13.750   1.4156   0.05256   0.04738  -0.0371   0.0036   1.0000
  14.000   1.4124   0.05598   0.05095  -0.0373   0.0035   1.0000
  14.250   1.4068   0.05989   0.05500  -0.0379   0.0034   1.0000
  14.500   1.4024   0.06385   0.05910  -0.0388   0.0033   1.0000
  14.750   1.3955   0.06840   0.06379  -0.0400   0.0032   1.0000
  15.000   1.3867   0.07345   0.06899  -0.0416   0.0032   1.0000
  15.250   1.3812   0.07817   0.07385  -0.0435   0.0031   1.0000
  15.500   1.3695   0.08407   0.07991  -0.0457   0.0031   1.0000
  15.750   1.3595   0.08985   0.08583  -0.0482   0.0031   1.0000
  16.000   1.3495   0.09582   0.09195  -0.0509   0.0031   1.0000
  16.250   1.3387   0.10210   0.09837  -0.0539   0.0030   1.0000
  16.500   1.3257   0.10894   0.10537  -0.0572   0.0030   1.0000
  16.750   1.3155   0.11547   0.11202  -0.0605   0.0030   1.0000
  17.000   1.3015   0.12298   0.11969  -0.0644   0.0030   1.0000
  17.250   1.2915   0.12992   0.12674  -0.0683   0.0030   1.0000
  17.500   1.2778   0.13781   0.13479  -0.0727   0.0030   1.0000
  17.750   1.2651   0.14578   0.14289  -0.0772   0.0030   1.0000
  18.000   1.2535   0.15368   0.15092  -0.0819   0.0030   1.0000
  18.250   1.2401   0.16230   0.15968  -0.0870   0.0030   1.0000
  18.500   1.2276   0.17094   0.16843  -0.0923   0.0030   1.0000
  18.750   1.2133   0.18039   0.17800  -0.0980   0.0030   1.0000
  19.000   1.1960   0.19140   0.18916  -0.1046   0.0031   1.0000
<< Back to GOE 371 AIRFOIL (goe371-il)

Polar data table (+)

Polar graphs


<< Back to GOE 371 AIRFOIL (goe371-il)