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BOEING-VERTOL VR-9 AIRFOIL (vr9-il) Xfoil prediction polar at RE=50,000 Ncrit=5


Details Polar file
Airfoil: BOEING-VERTOL VR-9 AIRFOIL (vr9-il)
Reynolds number: 50,000
Max Cl/Cd: 23.84 at α=3°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-vr9-il-50000-n5.txt
Download as CSV file: xf-vr9-il-50000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: BOEING-VERTOL VR-9 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 =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.000  -0.6656   0.11026   0.10372   0.0098   1.0000   0.1137
  -8.750  -0.6453   0.10523   0.09865   0.0135   1.0000   0.1233
  -8.500  -0.6592   0.10195   0.09550   0.0081   1.0000   0.1267
  -8.000  -0.6435   0.08710   0.08059  -0.0010   1.0000   0.0558
  -7.750  -0.6449   0.08070   0.07411  -0.0069   1.0000   0.0461
  -7.500  -0.6419   0.07611   0.06947  -0.0088   1.0000   0.0456
  -7.250  -0.6378   0.07146   0.06473  -0.0106   1.0000   0.0451
  -7.000  -0.6319   0.06687   0.06002  -0.0121   1.0000   0.0444
  -6.750  -0.6241   0.06230   0.05524  -0.0135   1.0000   0.0441
  -6.500  -0.6139   0.05781   0.05050  -0.0144   1.0000   0.0435
  -6.250  -0.6013   0.05337   0.04574  -0.0150   1.0000   0.0426
  -6.000  -0.5864   0.04903   0.04101  -0.0152   1.0000   0.0411
  -5.750  -0.5690   0.04481   0.03614  -0.0151   1.0000   0.0397
  -5.500  -0.5494   0.04092   0.03171  -0.0145   1.0000   0.0387
  -5.250  -0.5278   0.03739   0.02762  -0.0138   1.0000   0.0383
  -5.000  -0.5047   0.03420   0.02392  -0.0130   1.0000   0.0386
  -4.750  -0.4803   0.03140   0.02066  -0.0121   1.0000   0.0398
  -4.500  -0.4551   0.02936   0.01812  -0.0113   1.0000   0.0461
  -4.250  -0.4293   0.02696   0.01543  -0.0103   1.0000   0.0493
  -4.000  -0.4034   0.02493   0.01324  -0.0092   1.0000   0.0518
  -3.750  -0.3782   0.02332   0.01142  -0.0080   1.0000   0.0550
  -3.500  -0.3535   0.02202   0.00984  -0.0069   1.0000   0.0593
  -3.250  -0.3291   0.02090   0.00859  -0.0063   1.0000   0.0722
  -3.000  -0.3051   0.01973   0.00717  -0.0056   1.0000   0.0835
  -2.750  -0.2816   0.01845   0.00597  -0.0049   1.0000   0.1110
  -2.500  -0.2731   0.01537   0.00550  -0.0017   1.0000   0.5955
  -2.250  -0.2308   0.01466   0.00544  -0.0003   1.0000   0.9007
  -2.000  -0.1646   0.01448   0.00470  -0.0074   1.0000   0.9751
  -1.750  -0.1237   0.01427   0.00404  -0.0106   1.0000   1.0000
  -1.500  -0.1045   0.01412   0.00369  -0.0095   1.0000   1.0000
  -1.250  -0.0847   0.01401   0.00340  -0.0085   1.0000   1.0000
  -1.000  -0.0648   0.01393   0.00318  -0.0074   1.0000   1.0000
  -0.750  -0.0447   0.01388   0.00302  -0.0063   1.0000   1.0000
  -0.500  -0.0245   0.01385   0.00288  -0.0052   1.0000   1.0000
  -0.250  -0.0042   0.01385   0.00281  -0.0041   1.0000   1.0000
   0.000   0.0160   0.01386   0.00280  -0.0030   1.0000   1.0000
   0.250   0.0361   0.01390   0.00283  -0.0020   1.0000   1.0000
   0.500   0.0560   0.01396   0.00292  -0.0009   1.0000   1.0000
   0.750   0.0756   0.01406   0.00307   0.0001   1.0000   1.0000
   1.000   0.0948   0.01419   0.00326   0.0012   1.0000   1.0000
   1.250   0.1249   0.01436   0.00354  -0.0001   0.9881   1.0000
   1.500   0.1741   0.01455   0.00391  -0.0048   0.9580   1.0000
   1.750   0.2146   0.01473   0.00436  -0.0076   0.9334   1.0000
   2.000   0.2540   0.01491   0.00480  -0.0101   0.9085   1.0000
   2.250   0.2920   0.01510   0.00528  -0.0121   0.8829   1.0000
   2.500   0.3276   0.01530   0.00582  -0.0134   0.8554   1.0000
   2.750   0.3602   0.01541   0.00605  -0.0121   0.7745   1.0000
   3.000   0.3772   0.01582   0.00619  -0.0073   0.6301   1.0000
   3.250   0.3828   0.01822   0.00624  -0.0026   0.1801   1.0000
   3.500   0.4036   0.02010   0.00746  -0.0021   0.0800   1.0000
   3.750   0.4281   0.02123   0.00869  -0.0016   0.0663   1.0000
   4.000   0.4516   0.02249   0.01006  -0.0009   0.0606   1.0000
   4.250   0.4752   0.02374   0.01162   0.0002   0.0575   1.0000
   4.500   0.4993   0.02524   0.01338   0.0015   0.0547   1.0000
   4.750   0.5247   0.02710   0.01545   0.0027   0.0523   1.0000
   5.000   0.5494   0.02933   0.01800   0.0036   0.0460   1.0000
   5.250   0.5753   0.03169   0.02084   0.0046   0.0419   1.0000
   5.500   0.6000   0.03463   0.02428   0.0056   0.0408   1.0000
   5.750   0.6229   0.03797   0.02816   0.0066   0.0403   1.0000
   6.000   0.6434   0.04155   0.03230   0.0075   0.0401   1.0000
   6.250   0.6610   0.04492   0.03605   0.0080   0.0379   1.0000
   6.500   0.6736   0.04929   0.04051   0.0081   0.0350   1.0000
   6.750   0.6875   0.05291   0.04476   0.0085   0.0340   1.0000
   7.000   0.6980   0.05726   0.04949   0.0086   0.0340   1.0000
   7.250   0.7061   0.06189   0.05438   0.0084   0.0345   1.0000
   7.500   0.7128   0.06687   0.05949   0.0080   0.0351   1.0000
   7.750   0.7139   0.07108   0.06424   0.0063   0.0370   1.0000
   8.000   0.7090   0.07670   0.07006   0.0033   0.0388   1.0000
   8.250   0.7038   0.08212   0.07555   0.0001   0.0402   1.0000
   8.500   0.6969   0.08742   0.08082  -0.0037   0.0413   1.0000
   8.750   0.6932   0.09303   0.08637  -0.0080   0.0425   1.0000
   9.000   0.6919   0.09816   0.09145  -0.0110   0.0441   1.0000
   9.250   0.6937   0.10278   0.09602  -0.0127   0.0460   1.0000
   9.500   0.6918   0.10860   0.10181  -0.0172   0.0527   1.0000
   9.750   0.6923   0.11470   0.10787  -0.0205   0.0669   1.0000
  10.000   0.5902   0.10989   0.10348  -0.0132   0.0503   1.0000
  10.250   0.5835   0.11529   0.10882  -0.0163   0.0543   1.0000
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