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

GOE 373 AIRFOIL (goe373-il) Xfoil prediction polar at RE=200,000 Ncrit=5


Details Polar file
Airfoil: GOE 373 AIRFOIL (goe373-il)
Reynolds number: 200,000
Max Cl/Cd: 78.98 at α=5°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-goe373-il-200000-n5.txt
Download as CSV file: xf-goe373-il-200000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 373 AIRFOIL                                 
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.200 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -7.250  -0.2910   0.08898   0.08574  -0.0214   1.0000   0.0215
  -7.000  -0.2933   0.08699   0.08381  -0.0208   1.0000   0.0220
  -6.750  -0.2935   0.08486   0.08175  -0.0211   1.0000   0.0224
  -6.500  -0.2600   0.08030   0.07714  -0.0322   0.9906   0.0229
  -6.250  -0.2169   0.07527   0.07201  -0.0475   0.9794   0.0231
  -6.000  -0.2036   0.07103   0.06780  -0.0464   0.9744   0.0234
  -5.750  -0.1823   0.06738   0.06414  -0.0492   0.9649   0.0239
  -5.500  -0.1540   0.06361   0.06032  -0.0546   0.9547   0.0245
  -5.250  -0.1207   0.05968   0.05632  -0.0612   0.9455   0.0255
  -5.000  -0.0645   0.05524   0.05161  -0.0742   0.9339   0.0277
  -4.750  -0.0067   0.03268   0.02913  -0.0773   0.8831   0.0283
  -4.500   0.0130   0.02955   0.02594  -0.0785   0.8688   0.0288
  -4.250   0.0356   0.02667   0.02294  -0.0804   0.8532   0.0297
  -4.000   0.0606   0.02394   0.02006  -0.0826   0.8372   0.0309
  -3.750   0.1013   0.02157   0.01721  -0.0871   0.8217   0.0340
  -3.500   0.1257   0.01863   0.01403  -0.0885   0.8052   0.0344
  -3.250   0.1448   0.01629   0.01162  -0.0889   0.7885   0.0351
  -3.000   0.1673   0.01456   0.00975  -0.0894   0.7732   0.0362
  -2.750   0.1927   0.01300   0.00800  -0.0901   0.7601   0.0378
  -2.500   0.2265   0.01163   0.00610  -0.0911   0.7483   0.0425
  -2.250   0.2476   0.01012   0.00458  -0.0916   0.7358   0.0443
  -2.000   0.2729   0.00921   0.00351  -0.0919   0.7222   0.0482
  -1.750   0.3003   0.00818   0.00220  -0.0923   0.7089   0.0539
  -1.500   0.3267   0.00769   0.00149  -0.0922   0.6944   0.0606
  -1.250   0.3518   0.00682   0.00048  -0.0925   0.6807   0.0667
  -0.750   0.4192   0.01961   0.01259  -0.0977   0.6745   0.0831
   0.250   0.5309   0.01584   0.00757  -0.0949   0.6087   0.0499
   0.500   0.5577   0.01507   0.00657  -0.0942   0.5933   0.0444
   1.000   0.6102   0.01434   0.00554  -0.0930   0.5652   0.0396
   1.250   0.6362   0.01401   0.00509  -0.0924   0.5523   0.0383
   1.500   0.6620   0.01376   0.00473  -0.0918   0.5388   0.0373
   1.750   0.6877   0.01357   0.00447  -0.0912   0.5248   0.0366
   2.000   0.7132   0.01344   0.00427  -0.0906   0.5107   0.0363
   2.250   0.7388   0.01332   0.00413  -0.0901   0.4979   0.0364
   2.500   0.7643   0.01322   0.00403  -0.0896   0.4859   0.0376
   2.750   0.7897   0.01318   0.00398  -0.0892   0.4736   0.0400
   3.000   0.8153   0.01319   0.00396  -0.0887   0.4611   0.0404
   3.250   0.8410   0.01323   0.00396  -0.0882   0.4479   0.0405
   3.500   0.8666   0.01331   0.00399  -0.0877   0.4349   0.0406
   3.750   0.8921   0.01342   0.00407  -0.0872   0.4218   0.0409
   4.000   0.9175   0.01355   0.00416  -0.0867   0.4092   0.0414
   4.500   0.9680   0.01391   0.00444  -0.0857   0.3856   0.0439
   4.750   0.9933   0.01408   0.00466  -0.0852   0.3749   0.0499
   5.000   1.0228   0.01295   0.00503  -0.0860   0.3620   1.0000
   5.250   1.0465   0.01329   0.00527  -0.0853   0.3440   1.0000
   5.500   1.0698   0.01366   0.00554  -0.0845   0.3254   1.0000
   5.750   1.0932   0.01403   0.00585  -0.0838   0.3058   1.0000
   6.000   1.1165   0.01439   0.00616  -0.0831   0.2841   1.0000
   6.250   1.1386   0.01486   0.00649  -0.0823   0.2506   1.0000
   6.500   1.1559   0.01584   0.00702  -0.0810   0.1821   1.0000
   6.750   1.1749   0.01671   0.00766  -0.0798   0.1513   1.0000
   7.000   1.1953   0.01741   0.00825  -0.0788   0.1260   1.0000
   7.250   1.2088   0.01884   0.00920  -0.0771   0.0631   1.0000
   7.500   1.2213   0.02034   0.01042  -0.0751   0.0209   1.0000
   7.750   1.2401   0.02113   0.01128  -0.0737   0.0182   1.0000
   8.000   1.2592   0.02185   0.01213  -0.0724   0.0167   1.0000
   8.250   1.2777   0.02259   0.01301  -0.0711   0.0159   1.0000
   8.500   1.2945   0.02346   0.01403  -0.0696   0.0146   1.0000
   8.750   1.3098   0.02443   0.01515  -0.0679   0.0137   1.0000
   9.000   1.3232   0.02550   0.01641  -0.0659   0.0131   1.0000
   9.250   1.3338   0.02670   0.01777  -0.0637   0.0127   1.0000
   9.500   1.3389   0.02806   0.01929  -0.0607   0.0123   1.0000
   9.750   1.3390   0.02967   0.02105  -0.0574   0.0120   1.0000
  10.000   1.3359   0.03155   0.02307  -0.0541   0.0118   1.0000
  10.250   1.3361   0.03333   0.02495  -0.0515   0.0116   1.0000
  10.500   1.3388   0.03507   0.02681  -0.0496   0.0114   1.0000
  10.750   1.3418   0.03692   0.02880  -0.0480   0.0111   1.0000
  11.000   1.3435   0.03902   0.03103  -0.0467   0.0108   1.0000
  11.250   1.3445   0.04134   0.03347  -0.0456   0.0104   1.0000
  11.500   1.3444   0.04392   0.03616  -0.0448   0.0100   1.0000
  11.750   1.3436   0.04671   0.03905  -0.0442   0.0098   1.0000
  12.000   1.3430   0.04962   0.04206  -0.0438   0.0096   1.0000
  12.250   1.3427   0.05258   0.04512  -0.0435   0.0094   1.0000
  12.500   1.3431   0.05553   0.04816  -0.0432   0.0093   1.0000
  12.750   1.3441   0.05851   0.05125  -0.0430   0.0092   1.0000
  13.000   1.3454   0.06150   0.05433  -0.0429   0.0090   1.0000
  13.250   1.3472   0.06449   0.05741  -0.0427   0.0089   1.0000
  13.500   1.3500   0.06744   0.06045  -0.0426   0.0088   1.0000
  13.750   1.3528   0.07043   0.06354  -0.0424   0.0086   1.0000
  14.000   1.3559   0.07347   0.06667  -0.0424   0.0086   1.0000
  14.250   1.3596   0.07653   0.06983  -0.0423   0.0085   1.0000
  14.500   1.3634   0.07970   0.07310  -0.0422   0.0084   1.0000
  14.750   1.3648   0.08331   0.07686  -0.0424   0.0083   1.0000
  15.000   1.3639   0.08732   0.08103  -0.0430   0.0082   1.0000
  15.250   1.3624   0.09167   0.08553  -0.0437   0.0081   1.0000
  15.500   1.3522   0.09698   0.09104  -0.0460   0.0081   1.0000
  15.750   1.3418   0.10246   0.09673  -0.0488   0.0080   1.0000
  16.000   1.3310   0.10824   0.10271  -0.0519   0.0080   1.0000
  16.250   1.3194   0.11442   0.10910  -0.0555   0.0080   1.0000
<< Back to GOE 373 AIRFOIL (goe373-il)

Polar data table (+)

Polar graphs


<< Back to GOE 373 AIRFOIL (goe373-il)