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BOEING-VERTOL VR-8 AIRFOIL (vr8-il) Xfoil prediction polar at RE=100,000 Ncrit=9


Details Polar file
Airfoil: BOEING-VERTOL VR-8 AIRFOIL (vr8-il)
Reynolds number: 100,000
Max Cl/Cd: 44.26 at α=4.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-vr8-il-100000.txt
Download as CSV file: xf-vr8-il-100000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: BOEING-VERTOL VR-8 AIRFOIL                      
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.100 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.500  -0.5498   0.09196   0.08722  -0.0074   1.0000   0.0965
  -8.250  -0.5634   0.08795   0.08332  -0.0133   1.0000   0.0990
  -8.000  -0.5868   0.08388   0.07920  -0.0217   1.0000   0.1002
  -7.750  -0.5668   0.07890   0.07434  -0.0176   1.0000   0.1038
  -7.500  -0.5613   0.07551   0.07096  -0.0181   1.0000   0.1102
  -7.250  -0.5825   0.07229   0.06734  -0.0259   1.0000   0.1145
  -7.000  -0.5582   0.06749   0.06289  -0.0216   1.0000   0.1241
  -6.750  -0.5569   0.06324   0.05857  -0.0234   1.0000   0.1327
  -6.500  -0.5535   0.05965   0.05482  -0.0247   1.0000   0.1452
  -6.000  -0.5342   0.05310   0.04814  -0.0240   1.0000   0.1742
  -5.750  -0.5214   0.05011   0.04520  -0.0225   1.0000   0.1915
  -5.500  -0.5119   0.04739   0.04244  -0.0214   1.0000   0.2180
  -5.250  -0.5006   0.04491   0.04001  -0.0194   1.0000   0.2471
  -4.750  -0.4342   0.03195   0.02431  -0.0215   1.0000   0.0847
  -4.500  -0.4132   0.02865   0.02076  -0.0202   1.0000   0.0806
  -4.250  -0.3880   0.02679   0.01811  -0.0180   1.0000   0.0728
  -4.000  -0.3648   0.02411   0.01526  -0.0169   1.0000   0.0703
  -3.750  -0.3412   0.02210   0.01300  -0.0156   1.0000   0.0691
  -3.500  -0.3177   0.02091   0.01156  -0.0143   1.0000   0.0715
  -3.250  -0.2946   0.01962   0.01015  -0.0130   1.0000   0.0726
  -3.000  -0.2727   0.01814   0.00871  -0.0117   1.0000   0.0734
  -2.750  -0.2531   0.01699   0.00767  -0.0102   1.0000   0.0753
  -2.500  -0.2352   0.01622   0.00696  -0.0085   1.0000   0.0784
  -2.250  -0.2183   0.01568   0.00637  -0.0068   1.0000   0.0829
  -2.000  -0.2023   0.01521   0.00589  -0.0053   1.0000   0.0906
  -1.750  -0.1862   0.01480   0.00557  -0.0040   1.0000   0.1133
  -1.500  -0.0591   0.01165   0.00542  -0.0175   1.0000   1.0000
  -1.250  -0.0836   0.01194   0.00567  -0.0107   1.0000   1.0000
  -1.000  -0.0353   0.01218   0.00566  -0.0161   0.9879   1.0000
  -0.750   0.0207   0.01231   0.00560  -0.0225   0.9737   1.0000
  -0.500   0.0755   0.01238   0.00550  -0.0284   0.9599   1.0000
  -0.250   0.1512   0.01186   0.00487  -0.0368   0.9342   1.0000
   0.000   0.2000   0.01150   0.00438  -0.0398   0.8998   1.0000
   0.250   0.2316   0.01145   0.00422  -0.0403   0.8744   1.0000
   0.500   0.2590   0.01149   0.00416  -0.0400   0.8544   1.0000
   0.750   0.2843   0.01158   0.00417  -0.0393   0.8359   1.0000
   1.000   0.3084   0.01171   0.00425  -0.0385   0.8183   1.0000
   1.250   0.3323   0.01185   0.00435  -0.0376   0.8028   1.0000
   1.500   0.3560   0.01202   0.00448  -0.0367   0.7883   1.0000
   1.750   0.3781   0.01215   0.00455  -0.0352   0.7695   1.0000
   2.000   0.4000   0.01229   0.00461  -0.0335   0.7498   1.0000
   2.250   0.4227   0.01246   0.00480  -0.0323   0.7330   1.0000
   2.500   0.4449   0.01260   0.00493  -0.0308   0.7139   1.0000
   2.750   0.4672   0.01273   0.00502  -0.0293   0.6942   1.0000
   3.000   0.4892   0.01285   0.00516  -0.0277   0.6718   1.0000
   3.250   0.5115   0.01297   0.00527  -0.0262   0.6497   1.0000
   3.500   0.5337   0.01309   0.00547  -0.0247   0.6247   1.0000
   3.750   0.5555   0.01320   0.00560  -0.0230   0.5950   1.0000
   4.000   0.5773   0.01331   0.00578  -0.0214   0.5600   1.0000
   4.250   0.5958   0.01346   0.00576  -0.0190   0.4802   1.0000
   4.500   0.6054   0.01492   0.00600  -0.0158   0.2393   1.0000
   4.750   0.6170   0.01711   0.00725  -0.0137   0.0945   1.0000
   5.000   0.6364   0.01818   0.00828  -0.0123   0.0812   1.0000
   5.250   0.6566   0.01910   0.00931  -0.0108   0.0754   1.0000
   5.500   0.6756   0.02023   0.01044  -0.0093   0.0718   1.0000
   5.750   0.6950   0.02168   0.01184  -0.0078   0.0697   1.0000
   6.000   0.7183   0.02306   0.01327  -0.0067   0.0686   1.0000
   6.250   0.7436   0.02472   0.01506  -0.0058   0.0678   1.0000
   6.500   0.7684   0.02629   0.01682  -0.0049   0.0647   1.0000
   6.750   0.7938   0.02842   0.01918  -0.0040   0.0634   1.0000
   7.000   0.8174   0.03063   0.02158  -0.0031   0.0609   1.0000
   7.250   0.8387   0.03363   0.02475  -0.0024   0.0583   1.0000
   7.500   0.8590   0.03656   0.02833  -0.0005   0.0602   1.0000
   7.750   0.8719   0.04130   0.03391   0.0018   0.0650   1.0000
   9.750   0.7867   0.09825   0.09366  -0.0099   0.1666   1.0000
  10.000   0.7412   0.10455   0.09976  -0.0195   0.1576   1.0000
  10.250   0.7606   0.10813   0.10339  -0.0172   0.1534   1.0000
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