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

GOE 611 AIRFOIL (goe611-il) Xfoil prediction polar at RE=1,000,000 Ncrit=5


Details Polar file
Airfoil: GOE 611 AIRFOIL (goe611-il)
Reynolds number: 1,000,000
Max Cl/Cd: 107.87 at α=3°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-goe611-il-1000000-n5.txt
Download as CSV file: xf-goe611-il-1000000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 611 AIRFOIL                                 
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     1.000 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.250  -0.1484   0.08756   0.08509  -0.0721   0.7821   0.0054
  -9.000  -0.1461   0.08343   0.08097  -0.0744   0.7769   0.0056
  -8.750  -0.1459   0.07886   0.07639  -0.0771   0.7716   0.0058
  -8.500  -0.1500   0.07341   0.07094  -0.0808   0.7666   0.0060
  -8.250  -0.1571   0.06624   0.06377  -0.0885   0.7611   0.0062
  -8.000  -0.1464   0.06172   0.05921  -0.0943   0.7557   0.0063
  -7.000  -0.1390   0.01341   0.00813  -0.1132   0.7386   0.0111
  -6.750  -0.1101   0.01351   0.00825  -0.1134   0.7331   0.0117
  -6.500  -0.0811   0.01374   0.00850  -0.1136   0.7267   0.0121
  -6.250  -0.0523   0.01394   0.00871  -0.1138   0.7209   0.0125
  -6.000  -0.0263   0.01260   0.00713  -0.1138   0.7143   0.0139
  -5.750   0.0014   0.01224   0.00665  -0.1138   0.7074   0.0146
  -5.500   0.0305   0.01240   0.00682  -0.1141   0.6999   0.0149
  -5.250   0.0590   0.01253   0.00691  -0.1142   0.6919   0.0153
  -5.000   0.0876   0.01251   0.00685  -0.1144   0.6832   0.0158
  -4.750   0.1154   0.01218   0.00640  -0.1145   0.6746   0.0168
  -4.500   0.1429   0.01158   0.00563  -0.1145   0.6650   0.0180
  -4.250   0.1708   0.01135   0.00532  -0.1146   0.6551   0.0185
  -4.000   0.1990   0.01136   0.00528  -0.1147   0.6445   0.0189
  -3.750   0.2273   0.01141   0.00528  -0.1149   0.6332   0.0194
  -3.500   0.2553   0.01133   0.00512  -0.1150   0.6212   0.0200
  -3.250   0.2827   0.01114   0.00483  -0.1150   0.6052   0.0207
  -3.000   0.3102   0.01101   0.00457  -0.1149   0.5889   0.0216
  -2.750   0.3375   0.01089   0.00433  -0.1149   0.5722   0.0222
  -2.500   0.3651   0.01078   0.00412  -0.1149   0.5583   0.0226
  -2.250   0.3926   0.01071   0.00395  -0.1149   0.5436   0.0229
  -2.000   0.4196   0.01029   0.00347  -0.1149   0.5325   0.0238
  -1.750   0.4471   0.01015   0.00328  -0.1150   0.5229   0.0243
  -1.500   0.4750   0.01002   0.00311  -0.1151   0.5139   0.0246
  -1.250   0.5027   0.00992   0.00296  -0.1152   0.5055   0.0250
  -1.000   0.5306   0.00982   0.00282  -0.1153   0.4970   0.0254
  -0.750   0.5584   0.00976   0.00271  -0.1154   0.4890   0.0259
  -0.500   0.5863   0.00972   0.00262  -0.1156   0.4805   0.0267
  -0.250   0.6141   0.00967   0.00253  -0.1157   0.4725   0.0272
   0.000   0.6419   0.00965   0.00246  -0.1158   0.4636   0.0277
   0.250   0.6696   0.00965   0.00242  -0.1159   0.4555   0.0282
   0.500   0.6971   0.00967   0.00239  -0.1160   0.4461   0.0286
   0.750   0.7248   0.00970   0.00238  -0.1161   0.4369   0.0289
   1.000   0.7521   0.00976   0.00239  -0.1161   0.4272   0.0291
   1.250   0.7793   0.00980   0.00238  -0.1162   0.4163   0.0298
   1.500   0.8066   0.00986   0.00239  -0.1162   0.4060   0.0307
   1.750   0.8334   0.00996   0.00243  -0.1162   0.3949   0.0323
   2.000   0.8600   0.01008   0.00250  -0.1161   0.3832   0.0338
   2.250   0.8866   0.01020   0.00257  -0.1161   0.3709   0.0355
   2.750   0.9304   0.00880   0.00307  -0.1147   0.3439   0.8987
   3.000   0.9590   0.00889   0.00318  -0.1150   0.3303   1.0000
   3.250   0.9843   0.00913   0.00333  -0.1147   0.3165   1.0000
   3.500   1.0091   0.00940   0.00350  -0.1144   0.3007   1.0000
   3.750   1.0333   0.00970   0.00370  -0.1140   0.2832   1.0000
   4.000   1.0569   0.01004   0.00393  -0.1134   0.2648   1.0000
   4.250   1.0801   0.01040   0.00417  -0.1128   0.2459   1.0000
   4.500   1.1028   0.01078   0.00443  -0.1122   0.2263   1.0000
   4.750   1.1239   0.01125   0.00476  -0.1112   0.2030   1.0000
   5.000   1.1362   0.01227   0.00544  -0.1088   0.1405   1.0000
   5.250   1.1575   0.01267   0.00576  -0.1079   0.1280   1.0000
   5.500   1.1797   0.01300   0.00605  -0.1071   0.1209   1.0000
   5.750   1.2029   0.01325   0.00630  -0.1065   0.1186   1.0000
   6.000   1.2255   0.01351   0.00656  -0.1058   0.1167   1.0000
   6.250   1.2469   0.01383   0.00686  -0.1050   0.1140   1.0000
   6.500   1.2668   0.01419   0.00719  -0.1038   0.1098   1.0000
   6.750   1.2859   0.01450   0.00751  -0.1025   0.1066   1.0000
   7.000   1.3057   0.01479   0.00782  -0.1013   0.1051   1.0000
   7.250   1.3259   0.01508   0.00812  -0.1002   0.1041   1.0000
   7.500   1.3457   0.01540   0.00847  -0.0991   0.1023   1.0000
   7.750   1.3645   0.01578   0.00886  -0.0979   0.0994   1.0000
   8.000   1.3813   0.01627   0.00932  -0.0964   0.0947   1.0000
   8.250   1.3992   0.01672   0.00977  -0.0951   0.0902   1.0000
   8.500   1.4012   0.01805   0.01086  -0.0915   0.0541   1.0000
   8.750   1.3961   0.01989   0.01258  -0.0871   0.0162   1.0000
   9.000   1.4103   0.02065   0.01337  -0.0856   0.0130   1.0000
   9.250   1.4248   0.02141   0.01417  -0.0842   0.0115   1.0000
   9.500   1.4378   0.02231   0.01510  -0.0827   0.0099   1.0000
   9.750   1.4518   0.02315   0.01600  -0.0814   0.0092   1.0000
  10.000   1.4649   0.02409   0.01698  -0.0801   0.0085   1.0000
  10.250   1.4770   0.02513   0.01806  -0.0788   0.0078   1.0000
  10.500   1.4874   0.02634   0.01932  -0.0774   0.0071   1.0000
  10.750   1.4990   0.02747   0.02050  -0.0763   0.0068   1.0000
  11.000   1.5100   0.02869   0.02176  -0.0751   0.0063   1.0000
  11.250   1.5199   0.03002   0.02315  -0.0739   0.0059   1.0000
  11.500   1.5287   0.03148   0.02465  -0.0728   0.0055   1.0000
  11.750   1.5350   0.03316   0.02640  -0.0714   0.0052   1.0000
  12.000   1.5427   0.03476   0.02806  -0.0703   0.0050   1.0000
  12.250   1.5497   0.03646   0.02983  -0.0692   0.0048   1.0000
  12.500   1.5561   0.03827   0.03171  -0.0682   0.0046   1.0000
  12.750   1.5615   0.04020   0.03371  -0.0673   0.0044   1.0000
  13.000   1.5663   0.04224   0.03581  -0.0664   0.0042   1.0000
  13.250   1.5706   0.04439   0.03803  -0.0655   0.0040   1.0000
  13.500   1.5739   0.04671   0.04041  -0.0648   0.0038   1.0000
  13.750   1.5752   0.04929   0.04307  -0.0641   0.0037   1.0000
  14.000   1.5747   0.05213   0.04599  -0.0634   0.0035   1.0000
  14.250   1.5764   0.05480   0.04874  -0.0629   0.0034   1.0000
  14.500   1.5772   0.05763   0.05166  -0.0626   0.0033   1.0000
  14.750   1.5771   0.06064   0.05475  -0.0623   0.0032   1.0000
  15.000   1.5760   0.06380   0.05801  -0.0621   0.0032   1.0000
  15.250   1.5741   0.06710   0.06140  -0.0620   0.0031   1.0000
  15.500   1.5716   0.07054   0.06493  -0.0620   0.0030   1.0000
  15.750   1.5685   0.07411   0.06859  -0.0620   0.0029   1.0000
  16.000   1.5648   0.07780   0.07237  -0.0622   0.0028   1.0000
  16.250   1.5602   0.08165   0.07631  -0.0625   0.0028   1.0000
  16.500   1.5553   0.08555   0.08029  -0.0628   0.0027   1.0000
  16.750   1.5496   0.08962   0.08446  -0.0633   0.0027   1.0000
  17.000   1.5435   0.09381   0.08873  -0.0639   0.0026   1.0000
  17.250   1.5369   0.09811   0.09313  -0.0646   0.0026   1.0000
  17.500   1.5292   0.10260   0.09772  -0.0654   0.0025   1.0000
  17.750   1.5211   0.10719   0.10240  -0.0663   0.0025   1.0000
  18.000   1.5118   0.11203   0.10735  -0.0675   0.0024   1.0000
  18.250   1.5016   0.11711   0.11252  -0.0688   0.0024   1.0000
  18.500   1.4900   0.12249   0.11801  -0.0703   0.0023   1.0000
<< Back to GOE 611 AIRFOIL (goe611-il)

Polar data table (+)

Polar graphs


<< Back to GOE 611 AIRFOIL (goe611-il)