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

GOE 342 AIRFOIL (goe342-il) Xfoil prediction polar at RE=200,000 Ncrit=5


Details Polar file
Airfoil: GOE 342 AIRFOIL (goe342-il)
Reynolds number: 200,000
Max Cl/Cd: 85.95 at α=4°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-goe342-il-200000-n5.txt
Download as CSV file: xf-goe342-il-200000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 342 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
  ------ -------- --------- --------- -------- -------- --------
  -9.000  -0.2780   0.11078   0.10724  -0.0317   1.0000   0.0193
  -8.750  -0.2778   0.10891   0.10542  -0.0313   1.0000   0.0195
  -8.500  -0.2798   0.10729   0.10387  -0.0305   0.9997   0.0196
  -8.250  -0.2628   0.10387   0.10044  -0.0355   0.9942   0.0197
  -8.000  -0.2455   0.10062   0.09719  -0.0413   0.9871   0.0198
  -7.750  -0.2257   0.09733   0.09391  -0.0487   0.9779   0.0199
  -7.500  -0.2127   0.09213   0.08870  -0.0466   0.9766   0.0204
  -7.250  -0.1956   0.08869   0.08526  -0.0499   0.9690   0.0208
  -7.000  -0.1766   0.08532   0.08188  -0.0539   0.9611   0.0214
  -6.750  -0.1559   0.08191   0.07846  -0.0585   0.9535   0.0221
  -6.500  -0.1324   0.07846   0.07499  -0.0640   0.9459   0.0234
  -6.250  -0.0995   0.07543   0.07190  -0.0739   0.9353   0.0246
  -6.000  -0.0606   0.07203   0.06841  -0.0856   0.9253   0.0249
  -5.750  -0.0291   0.06806   0.06436  -0.0925   0.9185   0.0250
  -5.500  -0.0085   0.06374   0.06001  -0.0956   0.9100   0.0252
  -5.250   0.0025   0.06029   0.05657  -0.0942   0.9041   0.0257
  -5.000   0.0212   0.05756   0.05381  -0.0955   0.8962   0.0266
  -4.750   0.0489   0.05480   0.05098  -0.0994   0.8906   0.0289
  -4.500   0.1040   0.05179   0.04776  -0.1112   0.8834   0.0322
  -4.250   0.1476   0.04852   0.04429  -0.1184   0.8776   0.0325
  -3.750   0.1986   0.04119   0.03682  -0.1238   0.8640   0.0339
  -3.500   0.2254   0.03879   0.03434  -0.1258   0.8573   0.0350
  -3.250   0.2563   0.03647   0.03190  -0.1285   0.8497   0.0365
  -2.750   0.3444   0.03251   0.02725  -0.1370   0.8379   0.0433
  -2.500   0.3721   0.02891   0.02359  -0.1394   0.8322   0.0443
  -2.250   0.3998   0.02712   0.02172  -0.1407   0.8262   0.0456
  -2.000   0.4300   0.02555   0.02003  -0.1421   0.8196   0.0476
  -1.750   0.4492   0.00791   0.00224  -0.1355   0.8007   0.0552
  -1.250   0.5263   0.02139   0.01524  -0.1459   0.8021   0.0630
  -0.750   0.5941   0.01789   0.01099  -0.1473   0.7887   0.0416
  -0.500   0.6239   0.01675   0.00963  -0.1476   0.7796   0.0401
  -0.250   0.6534   0.01584   0.00846  -0.1477   0.7698   0.0392
   0.000   0.6825   0.01510   0.00753  -0.1478   0.7616   0.0388
   0.250   0.7114   0.01451   0.00676  -0.1478   0.7549   0.0388
   0.500   0.7396   0.01416   0.00629  -0.1477   0.7477   0.0409
   0.750   0.7678   0.01377   0.00579  -0.1475   0.7402   0.0414
   1.000   0.7958   0.01338   0.00534  -0.1474   0.7328   0.0411
   1.250   0.8235   0.01307   0.00499  -0.1473   0.7248   0.0409
   1.500   0.8511   0.01280   0.00472  -0.1471   0.7162   0.0410
   1.750   0.8786   0.01259   0.00449  -0.1469   0.7067   0.0412
   2.000   0.9058   0.01242   0.00435  -0.1467   0.6947   0.0416
   2.250   0.9327   0.01229   0.00425  -0.1463   0.6793   0.0423
   2.500   0.9591   0.01222   0.00415  -0.1459   0.6592   0.0430
   2.750   0.9853   0.01212   0.00402  -0.1455   0.6309   0.0459
   3.000   1.0101   0.01220   0.00394  -0.1447   0.5830   0.0500
   3.250   1.0321   0.01257   0.00395  -0.1434   0.5231   0.0517
   3.500   1.0539   0.01309   0.00415  -0.1422   0.4749   0.0537
   3.750   1.0762   0.01359   0.00444  -0.1412   0.4361   0.0591
   4.000   1.0950   0.01274   0.00488  -0.1397   0.4098   1.0000
   4.250   1.1188   0.01318   0.00518  -0.1390   0.3908   1.0000
   4.500   1.1430   0.01358   0.00548  -0.1383   0.3774   1.0000
   4.750   1.1670   0.01398   0.00581  -0.1377   0.3643   1.0000
   5.000   1.1907   0.01440   0.00618  -0.1370   0.3517   1.0000
   5.250   1.2141   0.01484   0.00655  -0.1363   0.3376   1.0000
   5.500   1.2378   0.01524   0.00692  -0.1356   0.3246   1.0000
   5.750   1.2619   0.01559   0.00731  -0.1350   0.3141   1.0000
   6.000   1.2854   0.01599   0.00772  -0.1343   0.3035   1.0000
   6.250   1.3085   0.01639   0.00813  -0.1336   0.2901   1.0000
   6.500   1.3315   0.01679   0.00854  -0.1328   0.2759   1.0000
   6.750   1.3546   0.01717   0.00899  -0.1320   0.2592   1.0000
   7.000   1.3768   0.01761   0.00944  -0.1312   0.2354   1.0000
   7.250   1.3948   0.01841   0.01001  -0.1298   0.1830   1.0000
   7.500   1.4077   0.01978   0.01098  -0.1277   0.1328   1.0000
   7.750   1.4240   0.02080   0.01190  -0.1261   0.1141   1.0000
   8.000   1.4411   0.02168   0.01279  -0.1246   0.1014   1.0000
   8.250   1.4597   0.02237   0.01354  -0.1232   0.0902   1.0000
   8.500   1.4774   0.02311   0.01431  -0.1218   0.0791   1.0000
   8.750   1.4940   0.02389   0.01511  -0.1202   0.0698   1.0000
   9.000   1.5091   0.02477   0.01597  -0.1184   0.0602   1.0000
   9.250   1.5228   0.02568   0.01692  -0.1164   0.0506   1.0000
   9.500   1.5342   0.02662   0.01791  -0.1140   0.0408   1.0000
   9.750   1.5430   0.02781   0.01907  -0.1113   0.0303   1.0000
  10.000   1.5501   0.02916   0.02041  -0.1085   0.0234   1.0000
  10.250   1.5573   0.03053   0.02186  -0.1059   0.0201   1.0000
  10.500   1.5625   0.03211   0.02351  -0.1032   0.0179   1.0000
  10.750   1.5665   0.03385   0.02539  -0.1005   0.0166   1.0000
  11.000   1.5716   0.03551   0.02724  -0.0981   0.0158   1.0000
  11.250   1.5753   0.03735   0.02925  -0.0957   0.0150   1.0000
  11.500   1.5778   0.03936   0.03142  -0.0935   0.0143   1.0000
  11.750   1.5793   0.04154   0.03375  -0.0914   0.0136   1.0000
  12.000   1.5789   0.04397   0.03632  -0.0894   0.0130   1.0000
  12.250   1.5749   0.04685   0.03935  -0.0875   0.0124   1.0000
  12.500   1.5672   0.05027   0.04290  -0.0856   0.0119   1.0000
  12.750   1.5670   0.05296   0.04577  -0.0843   0.0115   1.0000
  13.000   1.5644   0.05601   0.04900  -0.0831   0.0113   1.0000
  13.250   1.5606   0.05930   0.05250  -0.0821   0.0110   1.0000
  13.500   1.5557   0.06280   0.05617  -0.0812   0.0108   1.0000
  13.750   1.5500   0.06650   0.06005  -0.0806   0.0106   1.0000
  14.000   1.5436   0.07044   0.06417  -0.0802   0.0104   1.0000
  14.250   1.5367   0.07460   0.06851  -0.0801   0.0102   1.0000
  14.500   1.5292   0.07901   0.07310  -0.0804   0.0100   1.0000
  14.750   1.5213   0.08368   0.07794  -0.0810   0.0098   1.0000
  15.000   1.5129   0.08861   0.08305  -0.0819   0.0097   1.0000
  15.250   1.5040   0.09378   0.08840  -0.0832   0.0096   1.0000
  15.500   1.4948   0.09921   0.09400  -0.0849   0.0095   1.0000
  15.750   1.4854   0.10486   0.09982  -0.0869   0.0093   1.0000
  16.000   1.4757   0.11073   0.10586  -0.0892   0.0092   1.0000
  16.250   1.4659   0.11677   0.11207  -0.0919   0.0091   1.0000
  16.500   1.4557   0.12300   0.11845  -0.0948   0.0090   1.0000
  16.750   1.4455   0.12941   0.12502  -0.0981   0.0089   1.0000
  17.000   1.4354   0.13598   0.13173  -0.1016   0.0088   1.0000
  17.250   1.4249   0.14278   0.13868  -0.1054   0.0088   1.0000
  17.500   1.4144   0.14980   0.14584  -0.1095   0.0087   1.0000
  17.750   1.4033   0.15719   0.15338  -0.1141   0.0086   1.0000
  18.000   1.3920   0.16500   0.16135  -0.1191   0.0086   1.0000
  18.250   1.3792   0.17378   0.17031  -0.1249   0.0087   1.0000
  18.500   1.3618   0.18483   0.18157  -0.1325   0.0088   1.0000
<< Back to GOE 342 AIRFOIL (goe342-il)

Polar data table (+)

Polar graphs


<< Back to GOE 342 AIRFOIL (goe342-il)