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=500,000 Ncrit=9


Details Polar file
Airfoil: GOE 274 (DAIMLER V) AIRFOIL (goe274-il)
Reynolds number: 500,000
Max Cl/Cd: 98.98 at α=5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe274-il-500000.txt
Download as CSV file: xf-goe274-il-500000.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 =     0.500 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.500  -0.3367   0.10204   0.09985  -0.0214   1.0000   0.0183
  -8.250  -0.3420   0.10022   0.09807  -0.0197   1.0000   0.0185
  -8.000  -0.3494   0.09855   0.09643  -0.0176   1.0000   0.0188
  -7.750  -0.3474   0.09591   0.09382  -0.0184   0.9989   0.0193
  -7.500  -0.3312   0.09187   0.08978  -0.0238   0.9957   0.0202
  -7.250  -0.2969   0.08558   0.08345  -0.0413   0.9899   0.0208
  -7.000  -0.2634   0.07934   0.07715  -0.0538   0.9864   0.0208
  -6.750  -0.2479   0.07471   0.07253  -0.0549   0.9849   0.0211
  -6.500  -0.2252   0.07108   0.06889  -0.0577   0.9837   0.0213
  -6.250  -0.2034   0.06760   0.06538  -0.0615   0.9793   0.0217
  -6.000  -0.1741   0.06364   0.06139  -0.0676   0.9760   0.0223
  -5.750  -0.1385   0.05929   0.05697  -0.0752   0.9737   0.0235
  -5.500  -0.0726   0.05252   0.04989  -0.0912   0.9719   0.0252
  -5.250  -0.0326   0.04672   0.04389  -0.0984   0.9709   0.0254
  -5.000  -0.0128   0.04349   0.04068  -0.0998   0.9657   0.0257
  -4.750   0.0170   0.04081   0.03796  -0.1028   0.9628   0.0262
  -4.500   0.0514   0.03803   0.03509  -0.1065   0.9602   0.0270
  -4.250   0.0886   0.03503   0.03195  -0.1103   0.9578   0.0282
  -4.000   0.1313   0.03277   0.02924  -0.1125   0.9528   0.0306
  -3.750   0.1636   0.03066   0.02684  -0.1139   0.9474   0.0308
  -3.500   0.1915   0.02635   0.02250  -0.1163   0.9435   0.0315
  -3.250   0.2170   0.02472   0.02083  -0.1168   0.9364   0.0321
  -3.000   0.2456   0.02316   0.01916  -0.1176   0.9300   0.0330
  -2.750   0.2743   0.02169   0.01754  -0.1179   0.9220   0.0344
  -2.500   0.3074   0.02167   0.01710  -0.1172   0.9131   0.0374
  -2.250   0.3359   0.01883   0.01400  -0.1179   0.9013   0.0383
  -2.000   0.3629   0.01735   0.01249  -0.1181   0.8879   0.0392
  -1.750   0.3906   0.01638   0.01142  -0.1182   0.8730   0.0405
  -1.500   0.4189   0.01565   0.01053  -0.1181   0.8558   0.0430
  -1.250   0.4479   0.01653   0.01107  -0.1172   0.8351   0.0459
  -1.000   0.4760   0.01402   0.00842  -0.1178   0.8116   0.0484
  -0.750   0.5029   0.01346   0.00768  -0.1175   0.7793   0.0508
  -0.500   0.5292   0.01416   0.00804  -0.1165   0.7342   0.0559
   1.000   0.1235   0.05320   0.04814  -0.0084   0.5734   0.0460
   1.250   0.7146   0.01176   0.00400  -0.1136   0.4742   0.0474
   1.500   0.7422   0.01179   0.00397  -0.1135   0.4608   0.0457
   1.750   0.7701   0.01155   0.00371  -0.1135   0.4482   0.0452
   2.000   0.7978   0.01150   0.00360  -0.1135   0.4351   0.0455
   2.250   0.8253   0.01150   0.00355  -0.1135   0.4222   0.0459
   2.500   0.8533   0.01140   0.00342  -0.1136   0.4103   0.0484
   2.750   0.8811   0.01142   0.00342  -0.1137   0.3997   0.0498
   3.000   0.9086   0.01150   0.00344  -0.1136   0.3897   0.0516
   3.250   0.9358   0.01160   0.00349  -0.1135   0.3792   0.0550
   3.500   0.9633   0.01168   0.00357  -0.1134   0.3690   0.0666
   3.750   0.9826   0.01017   0.00384  -0.1120   0.3602   1.0000
   4.000   1.0095   0.01036   0.00396  -0.1119   0.3492   1.0000
   4.250   1.0364   0.01055   0.00409  -0.1117   0.3378   1.0000
   4.500   1.0630   0.01077   0.00426  -0.1116   0.3260   1.0000
   4.750   1.0894   0.01101   0.00444  -0.1114   0.3127   1.0000
   5.000   1.1155   0.01127   0.00463  -0.1112   0.2979   1.0000
   5.250   1.1413   0.01155   0.00485  -0.1109   0.2809   1.0000
   5.500   1.1663   0.01192   0.00509  -0.1105   0.2592   1.0000
   5.750   1.1913   0.01229   0.00535  -0.1102   0.2396   1.0000
   6.000   1.2162   0.01265   0.00564  -0.1098   0.2243   1.0000
   6.250   1.2408   0.01303   0.00596  -0.1093   0.2106   1.0000
   6.500   1.2652   0.01343   0.00629  -0.1089   0.1995   1.0000
   6.750   1.2902   0.01374   0.00660  -0.1085   0.1917   1.0000
   7.000   1.3144   0.01412   0.00695  -0.1080   0.1835   1.0000
   7.250   1.3387   0.01449   0.00730  -0.1076   0.1748   1.0000
   7.500   1.3627   0.01485   0.00766  -0.1071   0.1652   1.0000
   7.750   1.3862   0.01527   0.00805  -0.1065   0.1542   1.0000
   8.000   1.4090   0.01575   0.00844  -0.1058   0.1331   1.0000
   8.250   1.4211   0.01734   0.00954  -0.1036   0.0682   1.0000
   8.500   1.4324   0.01895   0.01088  -0.1011   0.0244   1.0000
   8.750   1.4510   0.01975   0.01174  -0.0996   0.0212   1.0000
   9.000   1.4704   0.02041   0.01251  -0.0983   0.0199   1.0000
   9.250   1.4884   0.02115   0.01335  -0.0967   0.0188   1.0000
   9.500   1.5042   0.02201   0.01431  -0.0949   0.0177   1.0000
   9.750   1.5155   0.02315   0.01555  -0.0923   0.0167   1.0000
  10.000   1.5220   0.02447   0.01701  -0.0890   0.0160   1.0000
  10.250   1.5314   0.02537   0.01800  -0.0861   0.0156   1.0000
  10.500   1.5389   0.02643   0.01917  -0.0830   0.0153   1.0000
  10.750   1.5447   0.02766   0.02051  -0.0799   0.0149   1.0000
  11.000   1.5490   0.02904   0.02200  -0.0768   0.0145   1.0000
  11.250   1.5519   0.03058   0.02365  -0.0738   0.0142   1.0000
  11.500   1.5533   0.03231   0.02549  -0.0709   0.0139   1.0000
  11.750   1.5532   0.03425   0.02754  -0.0682   0.0136   1.0000
  12.000   1.5517   0.03645   0.02984  -0.0657   0.0133   1.0000
  12.250   1.5474   0.03905   0.03255  -0.0634   0.0130   1.0000
  12.500   1.5402   0.04215   0.03576  -0.0615   0.0128   1.0000
  12.750   1.5295   0.04584   0.03956  -0.0599   0.0126   1.0000
  13.000   1.5164   0.05005   0.04388  -0.0586   0.0124   1.0000
  13.250   1.5090   0.05382   0.04776  -0.0579   0.0122   1.0000
  13.500   1.5082   0.05705   0.05112  -0.0578   0.0121   1.0000
  13.750   1.5061   0.06057   0.05476  -0.0580   0.0119   1.0000
  14.000   1.5029   0.06431   0.05862  -0.0582   0.0118   1.0000
  14.250   1.4995   0.06817   0.06259  -0.0586   0.0117   1.0000
  14.500   1.4959   0.07209   0.06663  -0.0591   0.0116   1.0000
  14.750   1.4924   0.07605   0.07069  -0.0596   0.0114   1.0000
  15.000   1.4889   0.08000   0.07476  -0.0602   0.0113   1.0000
  15.250   1.4856   0.08393   0.07880  -0.0607   0.0112   1.0000
  15.500   1.4821   0.08793   0.08290  -0.0614   0.0111   1.0000
  15.750   1.4789   0.09190   0.08698  -0.0621   0.0110   1.0000
  16.000   1.4756   0.09591   0.09110  -0.0628   0.0109   1.0000
  16.250   1.4720   0.10001   0.09531  -0.0637   0.0108   1.0000
  16.500   1.4682   0.10418   0.09959  -0.0647   0.0107   1.0000
  16.750   1.4641   0.10848   0.10401  -0.0659   0.0106   1.0000
  17.000   1.4595   0.11294   0.10860  -0.0673   0.0106   1.0000
  17.250   1.4546   0.11758   0.11336  -0.0691   0.0105   1.0000
  17.500   1.4490   0.12238   0.11829  -0.0711   0.0104   1.0000
  17.750   1.4434   0.12730   0.12333  -0.0733   0.0103   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)