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

BOEING-VERTOL VR-13 AIRFOIL (vr13-il) Xfoil prediction polar at RE=50,000 Ncrit=5


Details Polar file
Airfoil: BOEING-VERTOL VR-13 AIRFOIL (vr13-il)
Reynolds number: 50,000
Max Cl/Cd: 29.06 at α=6.25°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-vr13-il-50000-n5.txt
Download as CSV file: xf-vr13-il-50000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: BOEING-VERTOL VR-13 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
  ------ -------- --------- --------- -------- -------- --------
 -11.250  -0.5221   0.14003   0.13294   0.0086   1.0000   0.0943
 -11.000  -0.5296   0.13780   0.13079   0.0045   1.0000   0.0972
 -10.750  -0.5377   0.13532   0.12838   0.0002   1.0000   0.0979
 -10.500  -0.5126   0.12894   0.12200   0.0038   1.0000   0.1017
 -10.250  -0.5057   0.12527   0.11835   0.0031   1.0000   0.1050
 -10.000  -0.5041   0.12183   0.11495   0.0011   1.0000   0.1088
  -9.750  -0.5192   0.11945   0.11268  -0.0046   1.0000   0.1124
  -9.250  -0.5012   0.11039   0.10367  -0.0047   1.0000   0.1160
  -9.000  -0.4950   0.10656   0.09988  -0.0055   1.0000   0.1184
  -8.750  -0.4099   0.08803   0.08168  -0.0250   1.0000   0.0687
  -8.500  -0.4036   0.08387   0.07755  -0.0245   1.0000   0.0673
  -8.250  -0.4054   0.07937   0.07309  -0.0258   1.0000   0.0658
  -8.000  -0.5111   0.08218   0.07546  -0.0261   1.0000   0.0552
  -7.750  -0.5081   0.07806   0.07133  -0.0268   1.0000   0.0546
  -7.500  -0.5065   0.07389   0.06711  -0.0277   1.0000   0.0543
  -7.250  -0.5046   0.06975   0.06287  -0.0283   1.0000   0.0541
  -7.000  -0.5018   0.06574   0.05872  -0.0284   1.0000   0.0541
  -6.750  -0.4978   0.06190   0.05469  -0.0279   1.0000   0.0539
  -6.500  -0.4929   0.05825   0.05086  -0.0268   1.0000   0.0535
  -6.250  -0.4877   0.05479   0.04719  -0.0253   1.0000   0.0530
  -6.000  -0.4824   0.05153   0.04369  -0.0232   1.0000   0.0526
  -5.750  -0.4762   0.04846   0.04032  -0.0209   1.0000   0.0525
  -5.500  -0.4687   0.04562   0.03709  -0.0184   1.0000   0.0535
  -5.250  -0.4596   0.04304   0.03399  -0.0157   1.0000   0.0551
  -5.000  -0.4481   0.04069   0.03115  -0.0131   1.0000   0.0562
  -4.750  -0.4347   0.03803   0.02822  -0.0111   1.0000   0.0569
  -4.500  -0.4194   0.03588   0.02588  -0.0093   1.0000   0.0584
  -4.250  -0.4029   0.03425   0.02406  -0.0077   1.0000   0.0616
  -4.000  -0.3843   0.03256   0.02203  -0.0061   1.0000   0.0644
  -3.750  -0.3635   0.03087   0.01995  -0.0046   1.0000   0.0660
  -3.500  -0.3415   0.02947   0.01817  -0.0032   1.0000   0.0684
  -3.250  -0.3196   0.02811   0.01672  -0.0023   1.0000   0.0723
  -3.000  -0.2964   0.02702   0.01550  -0.0014   1.0000   0.0758
  -2.750  -0.2718   0.02605   0.01437  -0.0005   1.0000   0.0791
  -2.500  -0.2468   0.02528   0.01344   0.0001   1.0000   0.0848
  -2.250  -0.2217   0.02455   0.01267   0.0004   1.0000   0.0910
  -2.000  -0.1959   0.02398   0.01192   0.0008   0.9988   0.0968
  -1.750  -0.1606   0.02333   0.01122  -0.0010   0.9934   0.1083
  -1.500  -0.1259   0.02264   0.01066  -0.0027   0.9875   0.1355
  -1.250  -0.0414   0.01949   0.01059  -0.0121   0.9967   1.0000
  -1.000  -0.0040   0.01971   0.01042  -0.0145   0.9884   1.0000
  -0.750   0.0355   0.01994   0.01034  -0.0174   0.9792   1.0000
  -0.500   0.0763   0.02017   0.01032  -0.0204   0.9695   1.0000
  -0.250   0.1212   0.02039   0.01033  -0.0241   0.9579   1.0000
   0.000   0.1795   0.02043   0.01020  -0.0300   0.9403   1.0000
   0.250   0.2389   0.02026   0.00990  -0.0355   0.9180   1.0000
   0.500   0.2799   0.02012   0.00967  -0.0374   0.8962   1.0000
   0.750   0.3142   0.02005   0.00954  -0.0381   0.8784   1.0000
   1.000   0.3442   0.01997   0.00942  -0.0378   0.8598   1.0000
   1.250   0.3692   0.01992   0.00934  -0.0366   0.8388   1.0000
   1.500   0.3958   0.01977   0.00916  -0.0354   0.8175   1.0000
   1.750   0.4189   0.01967   0.00905  -0.0336   0.7932   1.0000
   2.000   0.4425   0.01953   0.00887  -0.0317   0.7671   1.0000
   2.250   0.4646   0.01942   0.00873  -0.0297   0.7362   1.0000
   2.500   0.4860   0.01934   0.00862  -0.0275   0.6986   1.0000
   2.750   0.5071   0.01926   0.00845  -0.0251   0.6514   1.0000
   3.000   0.5282   0.01926   0.00825  -0.0227   0.5881   1.0000
   3.250   0.5485   0.01949   0.00803  -0.0202   0.5171   1.0000
   3.500   0.5671   0.02006   0.00810  -0.0180   0.4562   1.0000
   3.750   0.5865   0.02077   0.00845  -0.0163   0.4120   1.0000
   4.000   0.6069   0.02150   0.00894  -0.0151   0.3799   1.0000
   4.250   0.6280   0.02220   0.00945  -0.0140   0.3552   1.0000
   4.500   0.6496   0.02292   0.00999  -0.0130   0.3363   1.0000
   4.750   0.6721   0.02360   0.01059  -0.0121   0.3195   1.0000
   5.000   0.6950   0.02430   0.01123  -0.0113   0.3052   1.0000
   5.250   0.7186   0.02499   0.01191  -0.0106   0.2926   1.0000
   5.500   0.7424   0.02569   0.01264  -0.0099   0.2810   1.0000
   5.750   0.7662   0.02643   0.01341  -0.0093   0.2705   1.0000
   6.000   0.7902   0.02720   0.01413  -0.0087   0.2615   1.0000
   6.250   0.8142   0.02802   0.01511  -0.0081   0.2529   1.0000
   6.500   0.8386   0.02887   0.01590  -0.0075   0.2460   1.0000
   6.750   0.8615   0.02978   0.01707  -0.0069   0.2376   1.0000
   7.000   0.8851   0.03067   0.01792  -0.0063   0.2308   1.0000
   7.250   0.9070   0.03174   0.01928  -0.0056   0.2235   1.0000
   7.500   0.9302   0.03277   0.02037  -0.0050   0.2178   1.0000
   7.750   0.9512   0.03398   0.02186  -0.0042   0.2113   1.0000
   8.000   0.9722   0.03505   0.02308  -0.0034   0.2045   1.0000
   8.250   0.9922   0.03627   0.02448  -0.0025   0.1981   1.0000
   8.500   1.0107   0.03745   0.02589  -0.0015   0.1907   1.0000
   8.750   1.0290   0.03857   0.02718  -0.0005   0.1836   1.0000
   9.000   1.0454   0.03982   0.02868   0.0006   0.1762   1.0000
   9.250   1.0634   0.04105   0.03004   0.0016   0.1702   1.0000
   9.500   1.0756   0.04294   0.03233   0.0030   0.1638   1.0000
   9.750   1.0956   0.04384   0.03322   0.0039   0.1583   1.0000
  10.000   1.1002   0.04625   0.03616   0.0058   0.1514   1.0000
  10.250   1.1157   0.04733   0.03732   0.0070   0.1455   1.0000
  10.500   1.1202   0.04987   0.04021   0.0087   0.1406   1.0000
  10.750   1.1199   0.05256   0.04326   0.0107   0.1354   1.0000
  11.000   1.1414   0.05290   0.04350   0.0115   0.1301   1.0000
  11.250   1.1202   0.05719   0.04831   0.0143   0.1263   1.0000
  11.500   1.1018   0.06079   0.05219   0.0166   0.1229   1.0000
  11.750   1.1009   0.06266   0.05415   0.0183   0.1190   1.0000
  12.000   1.1007   0.06476   0.05630   0.0195   0.1157   1.0000
  12.250   1.0599   0.07150   0.06328   0.0191   0.1153   1.0000
  12.500   1.0087   0.08138   0.07332   0.0149   0.1156   1.0000
  12.750   0.9310   0.09978   0.09169   0.0035   0.1161   1.0000
<< Back to BOEING-VERTOL VR-13 AIRFOIL (vr13-il)

Polar data table (+)

Polar graphs


<< Back to BOEING-VERTOL VR-13 AIRFOIL (vr13-il)