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GOE 419 AIRFOIL (goe419-il) Xfoil prediction polar at RE=200,000 Ncrit=9


Details Polar file
Airfoil: GOE 419 AIRFOIL (goe419-il)
Reynolds number: 200,000
Max Cl/Cd: 79.84 at α=2.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe419-il-200000.txt
Download as CSV file: xf-goe419-il-200000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 419 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
  ------ -------- --------- --------- -------- -------- --------
  -8.000  -0.3698   0.09578   0.09237  -0.0232   1.0000   0.0249
  -7.750  -0.3718   0.09398   0.09063  -0.0234   1.0000   0.0253
  -7.500  -0.3743   0.09224   0.08896  -0.0240   1.0000   0.0256
  -7.250  -0.3708   0.08995   0.08671  -0.0264   1.0000   0.0258
  -7.000  -0.3653   0.08739   0.08419  -0.0285   1.0000   0.0260
  -6.750  -0.3583   0.08476   0.08159  -0.0307   1.0000   0.0261
  -6.500  -0.3517   0.08206   0.07891  -0.0319   1.0000   0.0262
  -6.250  -0.3462   0.07940   0.07626  -0.0325   1.0000   0.0263
  -6.000  -0.3415   0.07683   0.07369  -0.0327   1.0000   0.0263
  -5.750  -0.3358   0.07416   0.07100  -0.0329   1.0000   0.0264
  -5.500  -0.3282   0.07133   0.06815  -0.0332   1.0000   0.0264
  -5.250  -0.3272   0.06574   0.06257  -0.0337   1.0000   0.0270
  -5.000  -0.3257   0.06192   0.05878  -0.0324   1.0000   0.0277
  -4.750  -0.3008   0.05770   0.05446  -0.0356   0.9969   0.0290
  -4.500  -0.2655   0.05342   0.05008  -0.0412   0.9930   0.0309
  -4.250  -0.2269   0.04909   0.04559  -0.0470   0.9883   0.0334
  -4.000  -0.1684   0.04532   0.04139  -0.0549   0.9843   0.0389
  -3.750  -0.1314   0.03950   0.03527  -0.0597   0.9802   0.0402
  -3.500  -0.0930   0.03835   0.03380  -0.0624   0.9746   0.0521
  -3.250  -0.0640   0.03281   0.02827  -0.0659   0.9714   0.0552
  -3.000  -0.0231   0.03082   0.02568  -0.0683   0.9655   0.0660
  -2.750   0.0090   0.02719   0.02207  -0.0708   0.9606   0.0706
  -2.500   0.0551   0.02333   0.01761  -0.0734   0.9574   0.0612
  -2.000   0.1336   0.01660   0.00954  -0.0753   0.9463   0.0468
  -1.750   0.1712   0.01514   0.00769  -0.0765   0.9406   0.0517
  -1.500   0.2089   0.01421   0.00664  -0.0783   0.9351   0.0654
  -1.250   0.2498   0.01296   0.00541  -0.0809   0.9320   0.0759
  -1.000   0.2839   0.01230   0.00474  -0.0821   0.9246   0.0887
  -0.750   0.3242   0.01139   0.00399  -0.0845   0.9199   0.1257
  -0.500   0.3842   0.00889   0.00370  -0.0917   0.9195   1.0000
  -0.250   0.4213   0.00878   0.00342  -0.0934   0.9100   1.0000
   0.000   0.4601   0.00867   0.00317  -0.0956   0.9008   1.0000
   0.250   0.4976   0.00859   0.00299  -0.0974   0.8904   1.0000
   0.500   0.5302   0.00859   0.00290  -0.0982   0.8777   1.0000
   0.750   0.5586   0.00865   0.00289  -0.0981   0.8631   1.0000
   1.000   0.5852   0.00873   0.00291  -0.0977   0.8477   1.0000
   1.250   0.6110   0.00882   0.00296  -0.0970   0.8322   1.0000
   1.500   0.6364   0.00891   0.00306  -0.0963   0.8164   1.0000
   1.750   0.6616   0.00900   0.00313  -0.0955   0.8007   1.0000
   2.000   0.6867   0.00909   0.00320  -0.0947   0.7849   1.0000
   2.250   0.7079   0.00910   0.00316  -0.0928   0.7544   1.0000
   2.500   0.7265   0.00910   0.00297  -0.0900   0.7017   1.0000
   2.750   0.7459   0.00935   0.00295  -0.0877   0.6465   1.0000
   3.000   0.7604   0.00990   0.00313  -0.0844   0.5550   1.0000
   3.250   0.7627   0.01139   0.00341  -0.0794   0.3405   1.0000
   3.500   0.7636   0.01411   0.00456  -0.0751   0.0400   1.0000
   3.750   0.7853   0.01477   0.00538  -0.0736   0.0334   1.0000
   4.000   0.8068   0.01544   0.00626  -0.0722   0.0321   1.0000
   4.250   0.8267   0.01626   0.00724  -0.0705   0.0309   1.0000
   4.500   0.8444   0.01727   0.00834  -0.0686   0.0264   1.0000
   4.750   0.8610   0.01852   0.00965  -0.0663   0.0258   1.0000
   5.000   0.8795   0.01989   0.01106  -0.0643   0.0266   1.0000
   5.250   0.9016   0.02155   0.01284  -0.0627   0.0290   1.0000
   5.500   0.9313   0.02430   0.01549  -0.0624   0.0337   1.0000
   5.750   0.9667   0.02585   0.01720  -0.0619   0.0458   1.0000
   6.000   1.0172   0.02755   0.01967  -0.0604   0.0861   1.0000
   7.750   1.0599   0.04272   0.03696  -0.0436   0.0327   1.0000
   8.000   1.0653   0.04195   0.03676  -0.0391   0.0292   1.0000
   8.250   1.0651   0.04561   0.04074  -0.0361   0.0274   1.0000
   8.500   1.0615   0.04949   0.04488  -0.0332   0.0262   1.0000
   8.750   1.0546   0.05333   0.04894  -0.0303   0.0254   1.0000
   9.000   1.0430   0.05704   0.05283  -0.0271   0.0248   1.0000
   9.250   1.0259   0.06032   0.05627  -0.0235   0.0246   1.0000
   9.500   1.0054   0.06385   0.05996  -0.0208   0.0245   1.0000
   9.750   0.9821   0.06794   0.06422  -0.0193   0.0247   1.0000
  10.000   0.9566   0.07285   0.06930  -0.0192   0.0251   1.0000
  10.250   0.9301   0.07870   0.07531  -0.0207   0.0258   1.0000
  10.500   0.9033   0.08556   0.08230  -0.0236   0.0266   1.0000
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