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GOE 456 AIRFOIL (goe456-il) Xfoil prediction polar at RE=500,000 Ncrit=9


Details Polar file
Airfoil: GOE 456 AIRFOIL (goe456-il)
Reynolds number: 500,000
Max Cl/Cd: 117.94 at α=3.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe456-il-500000.txt
Download as CSV file: xf-goe456-il-500000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 456 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
  ------ -------- --------- --------- -------- -------- --------
  -9.000  -0.3070   0.10451   0.10233  -0.0329   1.0000   0.0107
  -8.750  -0.3089   0.10233   0.10018  -0.0321   1.0000   0.0113
  -8.500  -0.3175   0.10117   0.09909  -0.0308   1.0000   0.0118
  -8.250  -0.3094   0.09808   0.09601  -0.0332   0.9987   0.0119
  -8.000  -0.2912   0.09408   0.09201  -0.0383   0.9960   0.0120
  -7.000  -0.2238   0.07386   0.07182  -0.0625   0.9770   0.0130
  -6.750  -0.2028   0.07119   0.06914  -0.0654   0.9738   0.0141
  -6.500  -0.1752   0.06708   0.06501  -0.0725   0.9709   0.0152
  -6.250  -0.1476   0.06244   0.06033  -0.0803   0.9658   0.0160
  -6.000  -0.1155   0.05743   0.05527  -0.0892   0.9607   0.0170
  -5.750  -0.0761   0.05177   0.04953  -0.0997   0.9567   0.0184
  -5.500  -0.0392   0.04742   0.04509  -0.1070   0.9489   0.0198
  -5.250  -0.0025   0.04193   0.03945  -0.1144   0.9432   0.0200
  -5.000   0.0255   0.03603   0.03331  -0.1191   0.9337   0.0201
  -4.750   0.0450   0.02690   0.02365  -0.1242   0.9241   0.0212
  -4.500   0.0683   0.02392   0.02047  -0.1252   0.9170   0.0224
  -4.250   0.0919   0.02180   0.01814  -0.1255   0.9092   0.0239
  -4.000   0.1179   0.01944   0.01542  -0.1257   0.9025   0.0257
  -3.750   0.1435   0.01503   0.01019  -0.1243   0.8942   0.0170
  -3.500   0.1707   0.01329   0.00805  -0.1239   0.8878   0.0171
  -3.250   0.1967   0.01268   0.00726  -0.1233   0.8796   0.0189
  -3.000   0.2238   0.01169   0.00606  -0.1228   0.8728   0.0194
  -2.750   0.2499   0.01073   0.00493  -0.1222   0.8646   0.0196
  -2.500   0.2758   0.00970   0.00374  -0.1216   0.8567   0.0211
  -2.250   0.3020   0.00900   0.00292  -0.1211   0.8486   0.0246
  -2.000   0.3284   0.00873   0.00258  -0.1206   0.8395   0.0302
  -1.750   0.3553   0.00830   0.00203  -0.1202   0.8312   0.0438
  -1.500   0.3810   0.00789   0.00194  -0.1198   0.8218   0.1577
  -1.250   0.4076   0.00785   0.00187  -0.1195   0.8116   0.1759
  -1.000   0.4343   0.00781   0.00180  -0.1192   0.8013   0.1927
  -0.750   0.4609   0.00774   0.00173  -0.1189   0.7910   0.2112
  -0.500   0.4873   0.00768   0.00166  -0.1185   0.7796   0.2293
  -0.250   0.5135   0.00760   0.00160  -0.1181   0.7669   0.2534
   0.000   0.5392   0.00750   0.00156  -0.1177   0.7524   0.2969
   0.250   0.5644   0.00734   0.00154  -0.1172   0.7363   0.3735
   0.500   0.6022   0.00600   0.00159  -0.1196   0.7198   1.0000
   0.750   0.6276   0.00611   0.00158  -0.1190   0.7038   1.0000
   1.000   0.6530   0.00623   0.00158  -0.1184   0.6891   1.0000
   1.250   0.6784   0.00635   0.00161  -0.1179   0.6751   1.0000
   1.500   0.7038   0.00649   0.00165  -0.1173   0.6616   1.0000
   1.750   0.7291   0.00664   0.00170  -0.1168   0.6473   1.0000
   2.000   0.7544   0.00678   0.00177  -0.1162   0.6340   1.0000
   2.250   0.7801   0.00692   0.00186  -0.1158   0.6235   1.0000
   2.500   0.8058   0.00707   0.00198  -0.1154   0.6147   1.0000
   2.750   0.8318   0.00719   0.00209  -0.1150   0.6055   1.0000
   3.000   0.8573   0.00733   0.00220  -0.1145   0.5943   1.0000
   3.250   0.8815   0.00751   0.00231  -0.1138   0.5752   1.0000
   3.500   0.9058   0.00768   0.00245  -0.1131   0.5552   1.0000
   3.750   0.9290   0.00790   0.00258  -0.1122   0.5274   1.0000
   4.000   0.9522   0.00815   0.00274  -0.1112   0.4968   1.0000
   4.250   0.9700   0.00874   0.00297  -0.1094   0.4170   1.0000
   4.500   0.9749   0.01059   0.00380  -0.1057   0.2328   1.0000
   4.750   0.9770   0.01289   0.00499  -0.1017   0.0300   1.0000
   5.000   0.9987   0.01342   0.00559  -0.1005   0.0198   1.0000
   5.250   1.0211   0.01387   0.00615  -0.0994   0.0184   1.0000
   5.500   1.0422   0.01442   0.00680  -0.0981   0.0165   1.0000
   5.750   1.0618   0.01509   0.00756  -0.0967   0.0147   1.0000
   6.000   1.0794   0.01592   0.00849  -0.0948   0.0137   1.0000
   6.250   1.0939   0.01698   0.00965  -0.0924   0.0129   1.0000
   6.500   1.1049   0.01831   0.01107  -0.0895   0.0124   1.0000
   6.750   1.1176   0.01951   0.01237  -0.0868   0.0125   1.0000
   7.000   1.1327   0.02061   0.01354  -0.0845   0.0128   1.0000
   7.250   1.1463   0.02243   0.01538  -0.0823   0.0123   1.0000
   7.500   1.1704   0.02492   0.01789  -0.0818   0.0119   1.0000
   7.750   1.1936   0.02594   0.01899  -0.0808   0.0124   1.0000
  14.750   0.8772   0.15163   0.14985  -0.0804   0.0115   1.0000
  15.000   0.8760   0.15656   0.15482  -0.0838   0.0117   1.0000
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