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

GOE 414 AIRFOIL (goe414-il) Xfoil prediction polar at RE=500,000 Ncrit=9


Details Polar file
Airfoil: GOE 414 AIRFOIL (goe414-il)
Reynolds number: 500,000
Max Cl/Cd: 100.03 at α=7.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe414-il-500000.txt
Download as CSV file: xf-goe414-il-500000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 414 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
  ------ -------- --------- --------- -------- -------- --------
 -10.750  -0.2197   0.10413   0.10189  -0.0700   0.9877   0.0283
 -10.500  -0.4440   0.03657   0.03369  -0.1270   0.9634   0.0184
 -10.250  -0.4419   0.03180   0.02842  -0.1298   0.9561   0.0188
 -10.000  -0.4352   0.02939   0.02564  -0.1290   0.9469   0.0191
  -9.750  -0.4275   0.02604   0.02185  -0.1285   0.9407   0.0196
  -9.500  -0.4135   0.02432   0.02001  -0.1273   0.9329   0.0202
  -9.250  -0.3926   0.02333   0.01891  -0.1268   0.9278   0.0208
  -9.000  -0.3720   0.02245   0.01792  -0.1262   0.9218   0.0216
  -8.750  -0.3503   0.02157   0.01687  -0.1256   0.9160   0.0226
  -8.500  -0.3270   0.02075   0.01582  -0.1250   0.9115   0.0236
  -8.250  -0.3063   0.01960   0.01441  -0.1242   0.9045   0.0244
  -8.000  -0.2864   0.01787   0.01255  -0.1234   0.8987   0.0256
  -7.750  -0.2626   0.01714   0.01174  -0.1229   0.8928   0.0267
  -7.500  -0.2382   0.01647   0.01096  -0.1225   0.8864   0.0279
  -7.250  -0.2122   0.01593   0.01028  -0.1222   0.8815   0.0294
  -7.000  -0.1872   0.01539   0.00959  -0.1217   0.8745   0.0304
  -6.750  -0.1645   0.01408   0.00816  -0.1210   0.8684   0.0323
  -6.500  -0.1392   0.01351   0.00754  -0.1207   0.8618   0.0338
  -6.250  -0.1133   0.01305   0.00699  -0.1203   0.8549   0.0356
  -6.000  -0.0869   0.01267   0.00649  -0.1200   0.8475   0.0371
  -5.750  -0.0618   0.01201   0.00573  -0.1195   0.8396   0.0390
  -5.500  -0.0360   0.01154   0.00522  -0.1192   0.8320   0.0414
  -5.250  -0.0093   0.01122   0.00481  -0.1189   0.8239   0.0440
  -5.000   0.0173   0.01093   0.00444  -0.1186   0.8150   0.0462
  -4.750   0.0441   0.01059   0.00406  -0.1183   0.8070   0.0515
  -4.500   0.0708   0.01033   0.00383  -0.1181   0.7977   0.0608
  -4.250   0.0983   0.01019   0.00368  -0.1180   0.7893   0.0785
  -4.000   0.1257   0.01008   0.00356  -0.1179   0.7797   0.0915
  -3.750   0.1529   0.00996   0.00342  -0.1178   0.7700   0.1005
  -3.500   0.1805   0.00995   0.00330  -0.1177   0.7601   0.1065
  -3.250   0.2072   0.00979   0.00313  -0.1175   0.7491   0.1130
  -3.000   0.2343   0.00972   0.00300  -0.1173   0.7379   0.1184
  -2.750   0.2614   0.00964   0.00285  -0.1171   0.7271   0.1242
  -2.500   0.2881   0.00955   0.00273  -0.1169   0.7164   0.1325
  -2.250   0.3150   0.00946   0.00262  -0.1168   0.7053   0.1419
  -2.000   0.3417   0.00936   0.00251  -0.1166   0.6949   0.1554
  -1.750   0.3680   0.00924   0.00242  -0.1164   0.6850   0.1812
  -1.500   0.3940   0.00905   0.00239  -0.1161   0.6750   0.2389
  -1.250   0.4204   0.00896   0.00240  -0.1160   0.6656   0.2916
  -1.000   0.4467   0.00893   0.00239  -0.1157   0.6565   0.3276
  -0.750   0.4734   0.00889   0.00240  -0.1156   0.6471   0.3577
  -0.500   0.4998   0.00887   0.00240  -0.1154   0.6383   0.3893
  -0.250   0.5259   0.00881   0.00242  -0.1151   0.6289   0.4291
   0.000   0.5510   0.00863   0.00245  -0.1148   0.6196   0.5034
   0.250   0.5698   0.00805   0.00254  -0.1130   0.6104   0.7404
   0.500   0.6233   0.00778   0.00272  -0.1181   0.5986   0.9800
   0.750   0.6718   0.00792   0.00276  -0.1228   0.5859   1.0000
   1.000   0.6953   0.00804   0.00279  -0.1220   0.5738   1.0000
   1.250   0.7184   0.00817   0.00283  -0.1211   0.5607   1.0000
   1.500   0.7413   0.00831   0.00288  -0.1202   0.5468   1.0000
   1.750   0.7640   0.00846   0.00294  -0.1192   0.5321   1.0000
   2.000   0.7865   0.00863   0.00301  -0.1182   0.5176   1.0000
   2.500   0.8323   0.00899   0.00320  -0.1164   0.4939   1.0000
   2.750   0.8556   0.00918   0.00332  -0.1156   0.4853   1.0000
   3.000   0.8791   0.00937   0.00344  -0.1148   0.4777   1.0000
   3.250   0.9031   0.00956   0.00359  -0.1142   0.4715   1.0000
   3.500   0.9273   0.00973   0.00373  -0.1136   0.4656   1.0000
   3.750   0.9511   0.00996   0.00389  -0.1129   0.4607   1.0000
   4.000   0.9761   0.01012   0.00406  -0.1125   0.4567   1.0000
   4.250   1.0010   0.01029   0.00422  -0.1121   0.4522   1.0000
   4.500   1.0252   0.01051   0.00440  -0.1115   0.4477   1.0000
   4.750   1.0496   0.01075   0.00461  -0.1111   0.4439   1.0000
   5.000   1.0748   0.01089   0.00480  -0.1107   0.4401   1.0000
   5.250   1.0987   0.01109   0.00499  -0.1101   0.4344   1.0000
   5.500   1.1217   0.01134   0.00520  -0.1094   0.4274   1.0000
   5.750   1.1454   0.01150   0.00540  -0.1088   0.4210   1.0000
   6.000   1.1681   0.01178   0.00563  -0.1080   0.4151   1.0000
   6.250   1.1914   0.01192   0.00583  -0.1073   0.4082   1.0000
   6.500   1.2139   0.01214   0.00606  -0.1066   0.4022   1.0000
   6.750   1.2361   0.01236   0.00630  -0.1057   0.3955   1.0000
   7.000   1.2571   0.01257   0.00651  -0.1046   0.3863   1.0000
   7.250   1.2764   0.01276   0.00673  -0.1032   0.3745   1.0000
   7.500   1.2953   0.01300   0.00698  -0.1017   0.3626   1.0000
   7.750   1.3126   0.01331   0.00726  -0.1000   0.3432   1.0000
   8.000   1.3186   0.01416   0.00780  -0.0965   0.2859   1.0000
   8.250   1.3104   0.01596   0.00912  -0.0912   0.2168   1.0000
   8.500   1.2993   0.01810   0.01076  -0.0858   0.1339   1.0000
   8.750   1.2811   0.02080   0.01300  -0.0799   0.0571   1.0000
   9.000   1.2779   0.02272   0.01476  -0.0762   0.0250   1.0000
   9.250   1.2891   0.02375   0.01587  -0.0744   0.0226   1.0000
   9.500   1.2982   0.02496   0.01714  -0.0725   0.0208   1.0000
   9.750   1.3056   0.02634   0.01861  -0.0706   0.0195   1.0000
  10.000   1.3150   0.02760   0.01996  -0.0689   0.0188   1.0000
  10.250   1.3226   0.02904   0.02148  -0.0673   0.0180   1.0000
  10.500   1.3284   0.03068   0.02319  -0.0656   0.0172   1.0000
  10.750   1.3318   0.03257   0.02517  -0.0638   0.0165   1.0000
  11.000   1.3321   0.03481   0.02750  -0.0621   0.0160   1.0000
  11.250   1.3279   0.03756   0.03036  -0.0602   0.0156   1.0000
  11.500   1.3197   0.04085   0.03377  -0.0585   0.0153   1.0000
  11.750   1.3202   0.04344   0.03646  -0.0574   0.0151   1.0000
  12.000   1.3212   0.04606   0.03917  -0.0565   0.0148   1.0000
  12.250   1.3209   0.04889   0.04210  -0.0557   0.0146   1.0000
  12.500   1.3194   0.05193   0.04523  -0.0550   0.0144   1.0000
  12.750   1.3176   0.05506   0.04846  -0.0544   0.0141   1.0000
  13.000   1.3156   0.05823   0.05171  -0.0539   0.0139   1.0000
  13.250   1.3138   0.06142   0.05499  -0.0534   0.0136   1.0000
  13.500   1.3123   0.06462   0.05826  -0.0531   0.0133   1.0000
  13.750   1.3113   0.06779   0.06150  -0.0528   0.0130   1.0000
  14.000   1.3100   0.07096   0.06472  -0.0525   0.0127   1.0000
  14.250   1.3085   0.07412   0.06793  -0.0521   0.0124   1.0000
  14.500   1.3076   0.07711   0.07095  -0.0516   0.0122   1.0000
  14.750   1.3087   0.07951   0.07336  -0.0505   0.0119   1.0000
  15.000   1.3179   0.08006   0.07387  -0.0478   0.0117   1.0000
  15.250   1.3227   0.08257   0.07648  -0.0477   0.0116   1.0000
  15.500   1.3283   0.08488   0.07889  -0.0474   0.0114   1.0000
  15.750   1.3347   0.08702   0.08111  -0.0470   0.0113   1.0000
<< Back to GOE 414 AIRFOIL (goe414-il)

Polar data table (+)

Polar graphs


<< Back to GOE 414 AIRFOIL (goe414-il)