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

GOE 802 A AIRFOIL (goe802a-il) Xfoil prediction polar at RE=200,000 Ncrit=5


Details Polar file
Airfoil: GOE 802 A AIRFOIL (goe802a-il)
Reynolds number: 200,000
Max Cl/Cd: 149.11 at α=8°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-goe802a-il-200000-n5.txt
Download as CSV file: xf-goe802a-il-200000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 802 A 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
  ------ -------- --------- --------- -------- -------- --------
  -8.500  -0.1596   0.09181   0.08833  -0.0573   0.9613   0.0272
  -8.250  -0.1449   0.08858   0.08509  -0.0615   0.9509   0.0280
  -8.000  -0.1324   0.08572   0.08221  -0.0661   0.9387   0.0283
  -7.750  -0.1132   0.08240   0.07884  -0.0729   0.9300   0.0285
  -7.500  -0.1005   0.07896   0.07540  -0.0755   0.9203   0.0287
  -7.250  -0.0896   0.07575   0.07218  -0.0752   0.9138   0.0291
  -7.000  -0.0748   0.07294   0.06935  -0.0772   0.9064   0.0298
  -6.750  -0.0583   0.07007   0.06644  -0.0801   0.8979   0.0305
  -6.500  -0.0388   0.06704   0.06335  -0.0835   0.8913   0.0314
  -6.250  -0.0180   0.06403   0.06029  -0.0873   0.8828   0.0325
  -6.000   0.0133   0.06086   0.05697  -0.0938   0.8752   0.0336
  -5.750   0.0450   0.05768   0.05360  -0.0996   0.8683   0.0339
  -5.500   0.0565   0.05452   0.05048  -0.0991   0.8607   0.0344
  -5.250   0.0760   0.05198   0.04787  -0.1002   0.8543   0.0353
  -5.000   0.1000   0.04949   0.04529  -0.1024   0.8469   0.0365
  -4.750   0.1265   0.04695   0.04263  -0.1048   0.8393   0.0379
  -4.250   0.1914   0.04225   0.03745  -0.1103   0.8245   0.0399
  -4.000   0.2117   0.03952   0.03468  -0.1110   0.8171   0.0404
  -3.750   0.2359   0.03747   0.03254  -0.1119   0.8091   0.0414
  -3.500   0.2629   0.03577   0.03069  -0.1129   0.8001   0.0442
  -3.250   0.2979   0.03485   0.02934  -0.1140   0.7915   0.0458
  -3.000   0.3205   0.03222   0.02669  -0.1146   0.7807   0.0464
  -2.750   0.3460   0.03047   0.02481  -0.1152   0.7702   0.0473
  -2.500   0.3730   0.02902   0.02319  -0.1156   0.7589   0.0488
  -2.250   0.4031   0.02840   0.02226  -0.1158   0.7477   0.0515
  -2.000   0.4328   0.02789   0.02137  -0.1158   0.7372   0.0518
  -1.750   0.4580   0.02556   0.01902  -0.1164   0.7254   0.0523
  -1.500   0.4842   0.02418   0.01753  -0.1167   0.7141   0.0539
  -1.250   0.5131   0.02423   0.01727  -0.1165   0.7012   0.0577
  -1.000   0.5407   0.02337   0.01619  -0.1165   0.6882   0.0577
  -0.500   0.5971   0.02000   0.01217  -0.1165   0.6639   0.0395
  -0.250   0.6241   0.01921   0.01122  -0.1165   0.6507   0.0392
   0.000   0.6509   0.01855   0.01037  -0.1164   0.6374   0.0390
   0.250   0.6777   0.01797   0.00958  -0.1162   0.6237   0.0390
   0.500   0.7044   0.01744   0.00885  -0.1160   0.6104   0.0398
   0.750   0.7307   0.01721   0.00857  -0.1160   0.5982   0.0414
   1.000   0.7570   0.01691   0.00814  -0.1158   0.5860   0.0425
   1.250   0.7834   0.01653   0.00763  -0.1156   0.5744   0.0425
   1.500   0.8094   0.01623   0.00720  -0.1154   0.5629   0.0428
   1.750   0.8354   0.01594   0.00683  -0.1151   0.5509   0.0434
   2.000   0.8611   0.01568   0.00650  -0.1149   0.5400   0.0445
   2.250   0.8867   0.01546   0.00621  -0.1146   0.5293   0.0460
   2.500   0.9127   0.01557   0.00632  -0.1144   0.5194   0.0491
   2.750   0.9382   0.01557   0.00629  -0.1142   0.5094   0.0529
   3.000   0.9638   0.01555   0.00627  -0.1139   0.4999   0.0566
   3.250   0.9891   0.01562   0.00630  -0.1136   0.4897   0.0628
   3.500   1.0147   0.01580   0.00645  -0.1133   0.4798   0.0729
   4.000   1.0650   0.01616   0.00668  -0.1126   0.4575   0.0921
   4.250   1.0889   0.01624   0.00672  -0.1121   0.4456   0.0939
   4.500   1.1131   0.01641   0.00688  -0.1117   0.4331   0.0944
   4.750   1.1371   0.01651   0.00701  -0.1113   0.4216   0.0949
   5.250   1.1843   0.01682   0.00735  -0.1104   0.3988   0.0964
   5.500   1.2073   0.01707   0.00759  -0.1098   0.3891   0.0967
   5.750   1.2313   0.01722   0.00782  -0.1094   0.3803   0.0971
   6.000   1.2542   0.01742   0.00805  -0.1089   0.3725   0.0978
   6.250   1.2777   0.01763   0.00832  -0.1084   0.3641   0.0986
   6.500   1.2996   0.01795   0.00865  -0.1077   0.3545   0.0993
   6.750   1.3222   0.01818   0.00893  -0.1071   0.3447   0.1010
   7.000   1.3429   0.01854   0.00925  -0.1062   0.3340   0.1030
   7.250   1.3647   0.01883   0.00959  -0.1055   0.3231   0.1049
   7.750   1.4056   0.01953   0.01030  -0.1036   0.3025   0.1104
   8.000   1.3703   0.00919   0.00032  -0.0938   0.2903   0.1145
   8.250   1.3856   0.00955   0.00068  -0.0921   0.2794   0.1165
   8.500   1.4631   0.02079   0.01157  -0.1002   0.2702   0.1175
   8.750   1.4792   0.02133   0.01210  -0.0986   0.2566   0.1200
   9.000   1.4921   0.02193   0.01267  -0.0965   0.2408   0.1244
   9.500   1.5108   0.02254   0.01430  -0.0915   0.2023   1.0000
   9.750   1.5190   0.02358   0.01527  -0.0891   0.1855   1.0000
  10.250   1.5370   0.02573   0.01738  -0.0849   0.1623   1.0000
  10.500   1.5451   0.02692   0.01856  -0.0829   0.1539   1.0000
  10.750   1.5551   0.02803   0.01971  -0.0812   0.1454   1.0000
  11.000   1.5616   0.02941   0.02109  -0.0793   0.1372   1.0000
  11.250   1.5709   0.03064   0.02239  -0.0777   0.1287   1.0000
  11.500   1.5763   0.03221   0.02397  -0.0760   0.1216   1.0000
  11.750   1.5853   0.03353   0.02537  -0.0746   0.1146   1.0000
  12.000   1.5899   0.03527   0.02712  -0.0730   0.1069   1.0000
  12.250   1.5961   0.03692   0.02883  -0.0717   0.0993   1.0000
  12.500   1.5987   0.03895   0.03088  -0.0703   0.0911   1.0000
  12.750   1.6023   0.04094   0.03292  -0.0691   0.0831   1.0000
  13.000   1.6032   0.04326   0.03527  -0.0679   0.0762   1.0000
  13.250   1.6038   0.04567   0.03773  -0.0668   0.0705   1.0000
  13.500   1.6019   0.04845   0.04054  -0.0659   0.0654   1.0000
  13.750   1.6008   0.05120   0.04336  -0.0651   0.0603   1.0000
  14.000   1.5961   0.05449   0.04669  -0.0645   0.0551   1.0000
  14.250   1.5945   0.05754   0.04983  -0.0641   0.0505   1.0000
  14.500   1.5900   0.06104   0.05341  -0.0638   0.0440   1.0000
  14.750   1.5811   0.06521   0.05759  -0.0638   0.0272   1.0000
  15.000   1.5657   0.07039   0.06275  -0.0641   0.0222   1.0000
  15.250   1.5533   0.07534   0.06775  -0.0646   0.0200   1.0000
  15.500   1.5442   0.07996   0.07249  -0.0652   0.0189   1.0000
  15.750   1.5355   0.08461   0.07728  -0.0660   0.0182   1.0000
  16.000   1.5267   0.08937   0.08219  -0.0669   0.0176   1.0000
  16.250   1.5170   0.09433   0.08729  -0.0680   0.0171   1.0000
  16.500   1.5069   0.09946   0.09257  -0.0692   0.0167   1.0000
  16.750   1.4963   0.10477   0.09802  -0.0707   0.0164   1.0000
  17.000   1.4855   0.11021   0.10361  -0.0724   0.0161   1.0000
  17.250   1.4742   0.11581   0.10937  -0.0743   0.0158   1.0000
<< Back to GOE 802 A AIRFOIL (goe802a-il)

Polar data table (+)

Polar graphs


<< Back to GOE 802 A AIRFOIL (goe802a-il)