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RG 15A-1.8/11.0 AIRFOIL (rg15a111-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: RG 15A-1.8/11.0 AIRFOIL (rg15a111-il)
Reynolds number: 50,000
Max Cl/Cd: 32.33 at α=6.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-rg15a111-il-50000.txt
Download as CSV file: xf-rg15a111-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: RG 15A-1.8/11.0 AIRFOIL                         
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.050 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.500  -0.4176   0.09731   0.09030  -0.0129   1.0000   0.3262
  -8.250  -0.4125   0.09402   0.08705  -0.0120   1.0000   0.3390
  -7.750  -0.5367   0.07353   0.06703  -0.0367   1.0000   0.1643
  -7.500  -0.5817   0.06431   0.05737  -0.0416   1.0000   0.1376
  -7.250  -0.5728   0.06008   0.05317  -0.0405   1.0000   0.1353
  -7.000  -0.5719   0.05549   0.04840  -0.0397   1.0000   0.1312
  -6.750  -0.5814   0.04953   0.04147  -0.0395   1.0000   0.1233
  -6.500  -0.5709   0.04563   0.03728  -0.0382   1.0000   0.1225
  -6.250  -0.5597   0.04202   0.03315  -0.0369   1.0000   0.1230
  -6.000  -0.5439   0.03937   0.03056  -0.0356   1.0000   0.1285
  -5.750  -0.5278   0.03663   0.02740  -0.0343   1.0000   0.1337
  -5.500  -0.5092   0.03374   0.02383  -0.0330   1.0000   0.1381
  -5.250  -0.4896   0.03163   0.02169  -0.0317   1.0000   0.1485
  -5.000  -0.4693   0.02965   0.01951  -0.0304   1.0000   0.1637
  -4.750  -0.4484   0.02776   0.01755  -0.0291   1.0000   0.1845
  -4.500  -0.4287   0.02624   0.01616  -0.0277   1.0000   0.2169
  -4.250  -0.4092   0.02482   0.01491  -0.0262   1.0000   0.2599
  -4.000  -0.3894   0.02343   0.01373  -0.0246   1.0000   0.3130
  -3.750  -0.3704   0.02200   0.01278  -0.0227   1.0000   0.3859
  -3.500  -0.3552   0.02069   0.01231  -0.0197   1.0000   0.4977
  -3.250  -0.3483   0.02042   0.01281  -0.0136   1.0000   0.6294
  -3.000  -0.3460   0.02069   0.01335  -0.0060   1.0000   0.7280
  -2.750  -0.3452   0.02091   0.01364   0.0018   1.0000   0.8026
  -2.500  -0.3363   0.02109   0.01375   0.0084   1.0000   0.8742
  -2.250  -0.0686   0.02222   0.01342  -0.0297   1.0000   1.0000
  -2.000  -0.0802   0.02187   0.01304  -0.0257   1.0000   1.0000
  -1.750  -0.0922   0.02156   0.01269  -0.0215   1.0000   1.0000
  -1.500  -0.1044   0.02126   0.01235  -0.0172   1.0000   1.0000
  -1.250  -0.1155   0.02096   0.01199  -0.0130   1.0000   1.0000
  -1.000  -0.1218   0.02075   0.01168  -0.0095   1.0000   1.0000
  -0.750  -0.1180   0.02072   0.01152  -0.0075   1.0000   1.0000
  -0.500  -0.1066   0.02088   0.01151  -0.0066   1.0000   1.0000
  -0.250  -0.0918   0.02117   0.01164  -0.0062   1.0000   1.0000
   0.000  -0.0755   0.02155   0.01188  -0.0060   1.0000   1.0000
   0.250  -0.0585   0.02201   0.01219  -0.0060   1.0000   1.0000
   0.500  -0.0414   0.02255   0.01261  -0.0060   1.0000   1.0000
   0.750  -0.0199   0.02324   0.01320  -0.0070   0.9981   1.0000
   1.000   0.0369   0.02468   0.01451  -0.0142   0.9804   1.0000
   1.250   0.0878   0.02581   0.01558  -0.0201   0.9619   1.0000
   1.500   0.1402   0.02689   0.01661  -0.0260   0.9440   1.0000
   1.750   0.1855   0.02774   0.01742  -0.0303   0.9249   1.0000
   2.000   0.2297   0.02848   0.01816  -0.0342   0.9052   1.0000
   2.250   0.2786   0.02915   0.01887  -0.0386   0.8862   1.0000
   2.500   0.3215   0.02971   0.01947  -0.0417   0.8666   1.0000
   2.750   0.3610   0.03019   0.02001  -0.0441   0.8460   1.0000
   3.000   0.4111   0.03042   0.02034  -0.0476   0.8269   1.0000
   3.250   0.4526   0.03064   0.02066  -0.0496   0.8066   1.0000
   3.500   0.4958   0.03068   0.02084  -0.0515   0.7858   1.0000
   3.750   0.5516   0.03018   0.02051  -0.0545   0.7670   1.0000
   4.000   0.5864   0.03016   0.02063  -0.0546   0.7449   1.0000
   4.250   0.6319   0.02959   0.02021  -0.0555   0.7240   1.0000
   4.500   0.6687   0.02925   0.02000  -0.0552   0.7011   1.0000
   4.750   0.7112   0.02853   0.01942  -0.0551   0.6781   1.0000
   5.000   0.7407   0.02840   0.01940  -0.0538   0.6521   1.0000
   5.250   0.7731   0.02811   0.01918  -0.0526   0.6257   1.0000
   5.500   0.8047   0.02786   0.01899  -0.0512   0.5980   1.0000
   5.750   0.8332   0.02779   0.01894  -0.0496   0.5685   1.0000
   6.000   0.8590   0.02789   0.01904  -0.0478   0.5378   1.0000
   6.250   0.8836   0.02812   0.01925  -0.0459   0.5058   1.0000
   6.500   0.9077   0.02843   0.01951  -0.0440   0.4730   1.0000
   6.750   0.9321   0.02883   0.01978  -0.0422   0.4399   1.0000
   7.000   0.9520   0.02956   0.02045  -0.0401   0.4067   1.0000
   7.250   0.9708   0.03045   0.02131  -0.0381   0.3739   1.0000
   7.500   0.9899   0.03148   0.02229  -0.0361   0.3424   1.0000
   7.750   1.0085   0.03262   0.02336  -0.0342   0.3121   1.0000
   8.000   1.0265   0.03399   0.02468  -0.0323   0.2842   1.0000
   8.250   1.0445   0.03544   0.02606  -0.0306   0.2582   1.0000
   8.500   1.0645   0.03693   0.02739  -0.0292   0.2341   1.0000
   8.750   1.0773   0.03899   0.02968  -0.0271   0.2143   1.0000
   9.000   1.0919   0.04097   0.03177  -0.0252   0.1960   1.0000
   9.250   1.1065   0.04347   0.03442  -0.0235   0.1816   1.0000
   9.500   1.1175   0.04606   0.03722  -0.0216   0.1690   1.0000
   9.750   1.1293   0.04895   0.04028  -0.0199   0.1588   1.0000
  10.000   1.1421   0.05182   0.04330  -0.0183   0.1498   1.0000
  10.250   1.1362   0.05559   0.04752  -0.0155   0.1448   1.0000
  10.500   1.1262   0.05933   0.05167  -0.0126   0.1407   1.0000
  10.750   1.1450   0.06275   0.05506  -0.0121   0.1338   1.0000
  11.000   1.1240   0.06713   0.05981  -0.0091   0.1332   1.0000
  11.250   1.1000   0.07149   0.06444  -0.0065   0.1331   1.0000
  11.500   1.0744   0.07615   0.06927  -0.0046   0.1335   1.0000
  11.750   1.0488   0.08146   0.07473  -0.0042   0.1340   1.0000
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