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

GOE 274 (DAIMLER V) AIRFOIL (goe274-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9


Details Polar file
Airfoil: GOE 274 (DAIMLER V) AIRFOIL (goe274-il)
Reynolds number: 1,000,000
Max Cl/Cd: 115.11 at α=4.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe274-il-1000000.txt
Download as CSV file: xf-goe274-il-1000000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 274 (DAIMLER V) AIRFOIL                     
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     1.000 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -11.500  -0.2713   0.12541   0.12387  -0.0244   1.0000   0.0117
 -11.250  -0.2712   0.12322   0.12170  -0.0237   1.0000   0.0120
 -11.000  -0.2704   0.12081   0.11929  -0.0233   0.9999   0.0123
 -10.750  -0.2618   0.11705   0.11553  -0.0255   0.9994   0.0129
 -10.500  -0.2537   0.11292   0.11140  -0.0280   0.9987   0.0136
 -10.250  -0.2479   0.10868   0.10716  -0.0308   0.9979   0.0138
 -10.000  -0.2430   0.10440   0.10287  -0.0338   0.9971   0.0139
  -9.750  -0.2388   0.09993   0.09840  -0.0368   0.9963   0.0139
  -9.500  -0.2288   0.09526   0.09373  -0.0380   0.9956   0.0140
  -9.250  -0.2176   0.09163   0.09010  -0.0397   0.9950   0.0142
  -9.000  -0.2066   0.08805   0.08652  -0.0419   0.9943   0.0144
  -8.750  -0.1971   0.08457   0.08304  -0.0438   0.9929   0.0147
  -8.500  -0.1877   0.08094   0.07941  -0.0460   0.9911   0.0150
  -8.250  -0.1777   0.07715   0.07562  -0.0485   0.9894   0.0156
  -8.000  -0.1674   0.07292   0.07140  -0.0518   0.9877   0.0167
  -7.750  -0.2697   0.08566   0.08408  -0.0422   0.9909   0.0150
  -6.500  -0.1419   0.05999   0.05828  -0.0836   0.9764   0.0174
  -6.250  -0.1182   0.05700   0.05526  -0.0870   0.9734   0.0176
  -6.000  -0.0925   0.05392   0.05214  -0.0910   0.9702   0.0179
  -5.750  -0.0702   0.05101   0.04918  -0.0939   0.9636   0.0184
  -5.500  -0.0429   0.04772   0.04581  -0.0977   0.9585   0.0197
  -5.250   0.0008   0.04226   0.04010  -0.1048   0.9524   0.0208
  -5.000   0.0280   0.03881   0.03649  -0.1070   0.9458   0.0208
  -4.750   0.0520   0.03373   0.03124  -0.1097   0.9396   0.0211
  -4.500   0.0748   0.03193   0.02939  -0.1104   0.9319   0.0214
  -4.250   0.1000   0.03020   0.02758  -0.1112   0.9253   0.0217
  -4.000   0.1268   0.02837   0.02564  -0.1121   0.9173   0.0222
  -3.750   0.1550   0.02648   0.02361  -0.1130   0.9083   0.0231
  -3.500   0.1894   0.02479   0.02159  -0.1130   0.8977   0.0251
  -3.250   0.2188   0.02307   0.01964  -0.1134   0.8849   0.0252
  -3.000   0.2480   0.02138   0.01769  -0.1138   0.8710   0.0253
  -2.750   0.2760   0.01786   0.01393  -0.1152   0.8557   0.0260
  -2.500   0.3033   0.01691   0.01284  -0.1154   0.8358   0.0264
  -2.250   0.3311   0.01608   0.01184  -0.1156   0.8133   0.0270
  -2.000   0.3590   0.01531   0.01086  -0.1157   0.7866   0.0278
  -1.750   0.3873   0.01471   0.00999  -0.1156   0.7541   0.0294
  -1.500   0.4151   0.01517   0.01009  -0.1148   0.7117   0.0306
  -1.250   0.4419   0.01490   0.00943  -0.1145   0.6514   0.0307
  -1.000   0.4690   0.01308   0.00712  -0.1148   0.5799   0.0317
  -0.750   0.4960   0.01265   0.00649  -0.1148   0.5377   0.0323
  -0.500   0.5237   0.01231   0.00600  -0.1149   0.5121   0.0332
  -0.250   0.5518   0.01207   0.00561  -0.1150   0.4914   0.0350
   0.000   0.5797   0.01277   0.00612  -0.1147   0.4723   0.0375
   0.250   0.6084   0.01141   0.00467  -0.1151   0.4577   0.0395
   0.500   0.6365   0.01112   0.00435  -0.1152   0.4445   0.0412
   0.750   0.6646   0.01100   0.00416  -0.1152   0.4309   0.0434
   1.000   0.6927   0.01076   0.00384  -0.1153   0.4174   0.0483
   1.250   0.7207   0.01055   0.00362  -0.1155   0.4049   0.0516
   1.500   0.7487   0.01051   0.00354  -0.1155   0.3940   0.0547
   1.750   0.7786   0.01001   0.00291  -0.1155   0.3847   0.0338
   2.000   0.8068   0.00997   0.00283  -0.1156   0.3749   0.0346
   2.250   0.8349   0.00996   0.00278  -0.1157   0.3647   0.0354
   2.500   0.8631   0.00996   0.00276  -0.1158   0.3553   0.0361
   2.750   0.8909   0.01002   0.00278  -0.1158   0.3451   0.0369
   3.000   0.9186   0.01009   0.00281  -0.1158   0.3341   0.0379
   3.250   0.9465   0.01011   0.00279  -0.1159   0.3236   0.0399
   3.500   0.9740   0.01021   0.00286  -0.1159   0.3126   0.0431
   3.750   1.0013   0.01035   0.00294  -0.1158   0.3004   0.0472
   4.000   1.0229   0.00889   0.00326  -0.1151   0.2880   1.0000
   4.250   1.0498   0.00912   0.00339  -0.1150   0.2728   1.0000
   4.500   1.0763   0.00938   0.00356  -0.1148   0.2544   1.0000
   4.750   1.1019   0.00975   0.00377  -0.1146   0.2289   1.0000
   5.000   1.1276   0.01009   0.00399  -0.1143   0.2079   1.0000
   5.250   1.1535   0.01040   0.00422  -0.1141   0.1942   1.0000
   5.500   1.1795   0.01069   0.00445  -0.1138   0.1840   1.0000
   5.750   1.2058   0.01093   0.00465  -0.1137   0.1781   1.0000
   6.000   1.2317   0.01121   0.00489  -0.1134   0.1702   1.0000
   6.250   1.2579   0.01144   0.00511  -0.1133   0.1647   1.0000
   6.500   1.2833   0.01175   0.00538  -0.1130   0.1573   1.0000
   6.750   1.3095   0.01195   0.00559  -0.1128   0.1521   1.0000
   7.000   1.3345   0.01228   0.00588  -0.1124   0.1426   1.0000
   7.250   1.3582   0.01275   0.00620  -0.1119   0.1180   1.0000
   7.500   1.3753   0.01395   0.00706  -0.1104   0.0685   1.0000
   7.750   1.3905   0.01535   0.00817  -0.1084   0.0197   1.0000
   8.000   1.4128   0.01590   0.00874  -0.1075   0.0167   1.0000
   8.250   1.4356   0.01637   0.00926  -0.1067   0.0156   1.0000
   8.500   1.4579   0.01684   0.00978  -0.1059   0.0146   1.0000
   8.750   1.4793   0.01739   0.01037  -0.1050   0.0136   1.0000
   9.000   1.4984   0.01814   0.01119  -0.1036   0.0126   1.0000
   9.250   1.5178   0.01881   0.01194  -0.1023   0.0121   1.0000
   9.500   1.5379   0.01936   0.01255  -0.1011   0.0117   1.0000
   9.750   1.5567   0.01998   0.01322  -0.0997   0.0112   1.0000
  10.000   1.5742   0.02064   0.01394  -0.0982   0.0107   1.0000
  10.250   1.5901   0.02136   0.01473  -0.0963   0.0103   1.0000
  10.500   1.6039   0.02216   0.01559  -0.0942   0.0100   1.0000
  10.750   1.6115   0.02323   0.01675  -0.0910   0.0096   1.0000
  11.000   1.6066   0.02490   0.01857  -0.0858   0.0092   1.0000
  11.250   1.6194   0.02560   0.01933  -0.0837   0.0090   1.0000
  11.500   1.6286   0.02655   0.02036  -0.0811   0.0088   1.0000
  11.750   1.6360   0.02768   0.02157  -0.0785   0.0086   1.0000
  12.000   1.6418   0.02895   0.02293  -0.0758   0.0084   1.0000
  12.250   1.6463   0.03039   0.02445  -0.0732   0.0082   1.0000
  12.500   1.6496   0.03198   0.02614  -0.0707   0.0080   1.0000
  12.750   1.6521   0.03374   0.02799  -0.0684   0.0078   1.0000
  13.000   1.6530   0.03574   0.03009  -0.0663   0.0077   1.0000
  13.250   1.6531   0.03795   0.03239  -0.0645   0.0075   1.0000
  13.500   1.6511   0.04053   0.03507  -0.0629   0.0074   1.0000
  13.750   1.6479   0.04341   0.03805  -0.0617   0.0073   1.0000
  14.000   1.6415   0.04690   0.04165  -0.0609   0.0072   1.0000
  14.250   1.6325   0.05099   0.04586  -0.0606   0.0071   1.0000
  14.500   1.6207   0.05575   0.05075  -0.0610   0.0070   1.0000
  14.750   1.6065   0.06123   0.05636  -0.0619   0.0070   1.0000
  15.000   1.5910   0.06722   0.06249  -0.0634   0.0069   1.0000
  15.250   1.5755   0.07330   0.06869  -0.0651   0.0069   1.0000
  15.500   1.5573   0.07971   0.07522  -0.0667   0.0068   1.0000
  15.750   1.5386   0.08602   0.08164  -0.0681   0.0067   1.0000
<< Back to GOE 274 (DAIMLER V) AIRFOIL (goe274-il)

Polar data table (+)

Polar graphs


<< Back to GOE 274 (DAIMLER V) AIRFOIL (goe274-il)