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GOE 458 AIRFOIL (goe458-il) Xfoil prediction polar at RE=200,000 Ncrit=9


Details Polar file
Airfoil: GOE 458 AIRFOIL (goe458-il)
Reynolds number: 200,000
Max Cl/Cd: 87.47 at α=4.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe458-il-200000.txt
Download as CSV file: xf-goe458-il-200000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 458 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 =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -7.500  -0.3326   0.09597   0.09290  -0.0261   1.0000   0.0306
  -7.250  -0.3510   0.09574   0.09276  -0.0230   1.0000   0.0316
  -7.000  -0.3672   0.09502   0.09211  -0.0197   1.0000   0.0316
  -6.750  -0.3413   0.09136   0.08846  -0.0327   0.9938   0.0328
  -6.500  -0.3106   0.08652   0.08359  -0.0428   0.9874   0.0331
  -6.250  -0.2954   0.07928   0.07636  -0.0473   0.9838   0.0341
  -6.000  -0.2792   0.07584   0.07291  -0.0470   0.9797   0.0359
  -5.750  -0.2540   0.07194   0.06896  -0.0520   0.9740   0.0381
  -5.500  -0.2177   0.06699   0.06395  -0.0614   0.9700   0.0415
  -5.250  -0.1609   0.06108   0.05780  -0.0795   0.9610   0.0457
  -5.000  -0.1397   0.05479   0.05152  -0.0835   0.9576   0.0477
  -4.750  -0.1169   0.05215   0.04886  -0.0848   0.9517   0.0505
  -4.500  -0.0600   0.04475   0.04098  -0.0990   0.9461   0.0598
  -4.250  -0.0327   0.04213   0.03846  -0.1004   0.9436   0.0629
  -4.000   0.0027   0.03769   0.03371  -0.1055   0.9364   0.0742
  -3.750   0.0415   0.03439   0.03016  -0.1098   0.9325   0.0870
  -3.500   0.0920   0.02249   0.01663  -0.1143   0.9305   0.0351
  -3.250   0.1250   0.02061   0.01454  -0.1157   0.9252   0.0392
  -3.000   0.1617   0.01827   0.01145  -0.1164   0.9204   0.0348
  -2.750   0.2012   0.01666   0.00955  -0.1181   0.9173   0.0343
  -2.500   0.2325   0.01520   0.00799  -0.1185   0.9112   0.0361
  -2.250   0.2663   0.01428   0.00702  -0.1193   0.9054   0.0420
  -2.000   0.3023   0.01332   0.00596  -0.1205   0.9010   0.0520
  -1.750   0.3300   0.01257   0.00555  -0.1201   0.8922   0.1790
  -1.500   0.3619   0.01250   0.00545  -0.1208   0.8854   0.2235
  -1.250   0.3917   0.01235   0.00529  -0.1210   0.8772   0.2530
  -1.000   0.4197   0.01215   0.00508  -0.1208   0.8679   0.2742
  -0.750   0.4511   0.01178   0.00471  -0.1212   0.8612   0.2937
  -0.500   0.4766   0.01151   0.00452  -0.1206   0.8503   0.3219
  -0.250   0.5022   0.01108   0.00435  -0.1200   0.8401   0.3916
   0.000   0.5471   0.00957   0.00407  -0.1232   0.8340   1.0000
   0.250   0.5728   0.00963   0.00398  -0.1225   0.8226   1.0000
   0.500   0.5990   0.00969   0.00392  -0.1220   0.8115   1.0000
   0.750   0.6256   0.00976   0.00387  -0.1215   0.8011   1.0000
   1.000   0.6525   0.00981   0.00380  -0.1211   0.7904   1.0000
   1.250   0.6790   0.00984   0.00374  -0.1205   0.7785   1.0000
   1.500   0.7049   0.00989   0.00369  -0.1199   0.7659   1.0000
   1.750   0.7304   0.00995   0.00369  -0.1192   0.7530   1.0000
   2.000   0.7558   0.01003   0.00371  -0.1185   0.7402   1.0000
   2.250   0.7812   0.01012   0.00375  -0.1179   0.7275   1.0000
   2.500   0.8065   0.01021   0.00382  -0.1172   0.7141   1.0000
   2.750   0.8319   0.01032   0.00390  -0.1165   0.7014   1.0000
   3.000   0.8575   0.01046   0.00403  -0.1160   0.6900   1.0000
   3.250   0.8832   0.01062   0.00418  -0.1154   0.6793   1.0000
   3.500   0.9092   0.01079   0.00433  -0.1150   0.6690   1.0000
   3.750   0.9345   0.01095   0.00455  -0.1143   0.6567   1.0000
   4.000   0.9564   0.01106   0.00460  -0.1128   0.6311   1.0000
   4.250   0.9784   0.01122   0.00474  -0.1114   0.6062   1.0000
   4.500   0.9989   0.01142   0.00484  -0.1096   0.5728   1.0000
   4.750   1.0171   0.01168   0.00499  -0.1075   0.5257   1.0000
   5.000   1.0298   0.01228   0.00523  -0.1043   0.4304   1.0000
   5.250   1.0248   0.01452   0.00618  -0.0990   0.2309   1.0000
   5.500   1.0241   0.01695   0.00748  -0.0948   0.0514   1.0000
   5.750   1.0397   0.01804   0.00852  -0.0925   0.0332   1.0000
   6.000   1.0583   0.01884   0.00952  -0.0907   0.0304   1.0000
   6.250   1.0758   0.01971   0.01057  -0.0888   0.0283   1.0000
   6.500   1.0905   0.02074   0.01175  -0.0865   0.0253   1.0000
   6.750   1.1017   0.02197   0.01310  -0.0837   0.0243   1.0000
   7.000   1.1108   0.02331   0.01457  -0.0807   0.0238   1.0000
   7.250   1.1198   0.02473   0.01605  -0.0775   0.0236   1.0000
   7.500   1.1331   0.02624   0.01761  -0.0750   0.0239   1.0000
   7.750   1.1555   0.02803   0.01947  -0.0738   0.0251   1.0000
   8.000   1.1925   0.03062   0.02218  -0.0746   0.0278   1.0000
   8.250   1.2227   0.03300   0.02472  -0.0745   0.0271   1.0000
  14.250   0.8462   0.16224   0.15951  -0.0815   0.0690   1.0000
  14.500   0.8586   0.16372   0.16102  -0.0796   0.0645   1.0000
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