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GOE 494 AIRFOIL (goe494-il) Xfoil prediction polar at RE=500,000 Ncrit=9


Details Polar file
Airfoil: GOE 494 AIRFOIL (goe494-il)
Reynolds number: 500,000
Max Cl/Cd: 135.38 at α=2.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe494-il-500000.txt
Download as CSV file: xf-goe494-il-500000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 494 AIRFOIL                                 
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.500 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.500  -0.3471   0.10113   0.09889  -0.0282   1.0000   0.0150
  -8.250  -0.3490   0.09910   0.09689  -0.0271   1.0000   0.0153
  -8.000  -0.3541   0.09760   0.09543  -0.0255   1.0000   0.0159
  -7.500  -0.2390   0.07429   0.07225  -0.0453   0.9887   0.0167
  -7.250  -0.2311   0.06788   0.06585  -0.0469   0.9866   0.0174
  -7.000  -0.2147   0.06405   0.06201  -0.0494   0.9845   0.0179
  -6.750  -0.2010   0.06025   0.05821  -0.0526   0.9807   0.0184
  -6.500  -0.1847   0.05618   0.05415  -0.0569   0.9758   0.0190
  -6.250  -0.1634   0.05162   0.04957  -0.0629   0.9729   0.0200
  -6.000  -0.1365   0.04660   0.04453  -0.0709   0.9709   0.0219
  -5.750  -0.1802   0.06190   0.05969  -0.0728   0.9731   0.0201
  -5.500  -0.1245   0.05622   0.05394  -0.0881   0.9710   0.0234
  -5.250  -0.0332   0.03017   0.02791  -0.1016   0.9598   0.0239
  -5.000   0.0094   0.01903   0.01654  -0.1180   0.9576   0.0247
  -4.750   0.0328   0.01668   0.01413  -0.1203   0.9522   0.0260
  -4.500   0.0632   0.01422   0.01155  -0.1241   0.9475   0.0281
  -4.250   0.1063   0.01169   0.00861  -0.1297   0.9446   0.0340
  -4.000   0.1413   0.00651   0.00282  -0.1365   0.9412   0.0363
  -3.750   0.1655   0.01935   0.01540  -0.1425   0.9480   0.0381
  -3.500   0.1997   0.01786   0.01368  -0.1444   0.9446   0.0430
  -3.250   0.2305   0.01611   0.01161  -0.1457   0.9393   0.0504
  -3.000   0.2599   0.01554   0.01092  -0.1461   0.9333   0.0568
  -2.750   0.2901   0.01442   0.00969  -0.1472   0.9284   0.0674
  -2.500   0.3219   0.01206   0.00673  -0.1465   0.9203   0.0452
  -2.250   0.3513   0.01083   0.00531  -0.1467   0.9135   0.0433
  -2.000   0.3787   0.01013   0.00448  -0.1464   0.9047   0.0429
  -1.750   0.4062   0.00955   0.00381  -0.1461   0.8961   0.0426
  -1.500   0.4341   0.00899   0.00316  -0.1459   0.8882   0.0414
  -1.250   0.4610   0.00858   0.00271  -0.1455   0.8792   0.0411
  -1.000   0.4889   0.00827   0.00234  -0.1454   0.8718   0.0418
  -0.750   0.5161   0.00803   0.00207  -0.1452   0.8633   0.0433
  -0.500   0.5437   0.00786   0.00187  -0.1450   0.8557   0.0446
  -0.250   0.5711   0.00774   0.00171  -0.1448   0.8481   0.0456
   0.000   0.5985   0.00766   0.00160  -0.1446   0.8405   0.0464
   0.250   0.6260   0.00759   0.00149  -0.1444   0.8335   0.0502
   0.500   0.6531   0.00754   0.00146  -0.1441   0.8246   0.0560
   0.750   0.6802   0.00718   0.00147  -0.1440   0.8154   0.2160
   1.000   0.7069   0.00697   0.00156  -0.1438   0.8060   0.3495
   1.250   0.7334   0.00676   0.00169  -0.1436   0.7962   0.4721
   1.500   0.7571   0.00583   0.00176  -0.1426   0.7839   1.0000
   1.750   0.7826   0.00595   0.00176  -0.1418   0.7655   1.0000
   2.000   0.8081   0.00609   0.00181  -0.1411   0.7474   1.0000
   2.250   0.8336   0.00622   0.00188  -0.1404   0.7289   1.0000
   2.500   0.8591   0.00637   0.00199  -0.1398   0.7104   1.0000
   2.750   0.8840   0.00653   0.00209  -0.1389   0.6849   1.0000
   3.000   0.9079   0.00672   0.00217  -0.1379   0.6451   1.0000
   3.250   0.9253   0.00741   0.00234  -0.1357   0.5298   1.0000
   3.500   0.9345   0.00933   0.00308  -0.1326   0.2971   1.0000
   3.750   0.9436   0.01165   0.00413  -0.1299   0.0523   1.0000
   4.000   0.9664   0.01232   0.00480  -0.1288   0.0336   1.0000
   4.250   0.9897   0.01291   0.00543  -0.1279   0.0275   1.0000
   4.500   1.0113   0.01373   0.00633  -0.1266   0.0233   1.0000
   4.750   1.0342   0.01435   0.00700  -0.1256   0.0213   1.0000
   5.000   1.0561   0.01509   0.00781  -0.1244   0.0194   1.0000
   5.250   1.0772   0.01596   0.00872  -0.1231   0.0179   1.0000
   5.500   1.0936   0.01774   0.01056  -0.1210   0.0161   1.0000
   5.750   1.1148   0.01902   0.01191  -0.1197   0.0152   1.0000
   6.000   1.1378   0.02015   0.01312  -0.1187   0.0145   1.0000
   6.250   1.1615   0.02162   0.01468  -0.1177   0.0139   1.0000
   6.500   1.1862   0.02363   0.01686  -0.1167   0.0135   1.0000
   6.750   1.2117   0.02674   0.02021  -0.1157   0.0139   1.0000
   7.000   1.2396   0.03262   0.02672  -0.1131   0.0205   1.0000
   7.250   1.2590   0.03449   0.02880  -0.1117   0.0189   1.0000
   7.500   1.2768   0.03687   0.03137  -0.1101   0.0180   1.0000
   7.750   1.2939   0.03954   0.03415  -0.1089   0.0172   1.0000
   8.000   1.3034   0.04473   0.03958  -0.1071   0.0163   1.0000
   8.250   1.2927   0.05440   0.04986  -0.1029   0.0157   1.0000
   8.500   1.2962   0.05801   0.05381  -0.1000   0.0156   1.0000
   8.750   1.2971   0.06130   0.05741  -0.0969   0.0155   1.0000
   9.000   1.3083   0.06055   0.05700  -0.0936   0.0146   1.0000
   9.250   1.3090   0.06344   0.06019  -0.0903   0.0136   1.0000
   9.500   1.3012   0.06725   0.06423  -0.0871   0.0131   1.0000
   9.750   1.2862   0.07073   0.06790  -0.0832   0.0129   1.0000
  10.000   1.2666   0.07417   0.07151  -0.0796   0.0129   1.0000
  10.250   1.2456   0.07825   0.07573  -0.0773   0.0129   1.0000
  10.500   1.2234   0.08306   0.08070  -0.0766   0.0129   1.0000
  10.750   1.2006   0.08873   0.08651  -0.0775   0.0130   1.0000
  11.000   1.1767   0.09551   0.09342  -0.0803   0.0132   1.0000
  11.250   1.1531   0.10362   0.10161  -0.0850   0.0135   1.0000
  11.500   1.1312   0.11304   0.11112  -0.0915   0.0141   1.0000
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