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BOEING-VERTOL VR-8 AIRFOIL WITH TAB (vr8b-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: BOEING-VERTOL VR-8 AIRFOIL WITH TAB (vr8b-il)
Reynolds number: 50,000
Max Cl/Cd: 32.49 at α=5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-vr8b-il-50000.txt
Download as CSV file: xf-vr8b-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: BOEING-VERTOL VR-8 AIRFOIL WITH TAB             
  
 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
  ------ -------- --------- --------- -------- -------- --------
 -10.500  -0.6870   0.14621   0.13987   0.0441   1.0000   0.1671
 -10.250  -0.6654   0.13976   0.13340   0.0460   1.0000   0.1758
 -10.000  -0.6802   0.13831   0.13204   0.0416   1.0000   0.1814
  -9.750  -0.6596   0.13231   0.12602   0.0435   1.0000   0.1921
  -9.500  -0.6573   0.12834   0.12205   0.0422   1.0000   0.2001
  -9.250  -0.6662   0.12608   0.11986   0.0393   1.0000   0.2098
  -9.000  -0.6568   0.12169   0.11550   0.0397   1.0000   0.2233
  -8.750  -0.6514   0.11778   0.11163   0.0396   1.0000   0.2374
  -8.500  -0.6490   0.11418   0.10807   0.0391   1.0000   0.2521
  -8.250  -0.6498   0.11087   0.10483   0.0383   1.0000   0.2672
  -8.000  -0.6351   0.10615   0.10012   0.0398   1.0000   0.2853
  -7.750  -0.6252   0.10245   0.09644   0.0407   1.0000   0.3062
  -7.500  -0.6391   0.10033   0.09444   0.0391   1.0000   0.3250
  -7.250  -0.6203   0.09606   0.09017   0.0417   1.0000   0.3525
  -7.000  -0.6093   0.09247   0.08660   0.0436   1.0000   0.3808
  -6.750  -0.6082   0.08973   0.08393   0.0451   1.0000   0.4109
  -6.500  -0.5786   0.08525   0.07937   0.0487   1.0000   0.4500
  -6.250  -0.5652   0.08219   0.07633   0.0519   1.0000   0.4948
  -6.000  -0.5434   0.07825   0.07239   0.0543   1.0000   0.5348
  -5.250  -0.5147   0.04475   0.03672  -0.0181   1.0000   0.1498
  -5.000  -0.4842   0.04034   0.03161  -0.0195   1.0000   0.1356
  -4.750  -0.4516   0.03697   0.02722  -0.0203   1.0000   0.1245
  -4.500  -0.4225   0.03331   0.02330  -0.0206   1.0000   0.1198
  -4.250  -0.3912   0.03043   0.01991  -0.0206   1.0000   0.1156
  -4.000  -0.3603   0.02806   0.01713  -0.0204   1.0000   0.1149
  -3.750  -0.3306   0.02618   0.01502  -0.0203   1.0000   0.1211
  -3.500  -0.3005   0.02452   0.01308  -0.0197   1.0000   0.1246
  -3.250  -0.2726   0.02273   0.01136  -0.0188   1.0000   0.1296
  -3.000  -0.2445   0.02144   0.00996  -0.0181   1.0000   0.1383
  -2.750  -0.2179   0.02006   0.00866  -0.0178   1.0000   0.1547
  -2.500  -0.1922   0.01770   0.00719  -0.0181   1.0000   0.2513
  -2.250  -0.1192   0.01566   0.00719  -0.0167   1.0000   1.0000
  -2.000  -0.0994   0.01534   0.00659  -0.0168   1.0000   1.0000
  -1.750  -0.0807   0.01512   0.00615  -0.0166   1.0000   1.0000
  -1.500  -0.0640   0.01501   0.00584  -0.0159   1.0000   1.0000
  -1.250  -0.0506   0.01500   0.00567  -0.0147   1.0000   1.0000
  -1.000  -0.0398   0.01512   0.00563  -0.0131   1.0000   1.0000
  -0.750  -0.0272   0.01535   0.00566  -0.0118   1.0000   1.0000
  -0.500  -0.0122   0.01565   0.00579  -0.0112   1.0000   1.0000
  -0.250   0.0050   0.01601   0.00600  -0.0110   1.0000   1.0000
   0.000   0.0237   0.01643   0.00628  -0.0112   1.0000   1.0000
   0.250   0.0432   0.01689   0.00663  -0.0115   1.0000   1.0000
   0.500   0.0634   0.01739   0.00702  -0.0120   1.0000   1.0000
   0.750   0.0840   0.01793   0.00749  -0.0126   1.0000   1.0000
   1.000   0.1048   0.01852   0.00801  -0.0133   1.0000   1.0000
   1.250   0.1532   0.01924   0.00872  -0.0191   0.9885   1.0000
   1.500   0.2079   0.02003   0.00954  -0.0259   0.9740   1.0000
   1.750   0.2859   0.02058   0.01022  -0.0358   0.9413   1.0000
   2.000   0.3472   0.02094   0.01077  -0.0422   0.9192   1.0000
   2.250   0.3839   0.02144   0.01139  -0.0443   0.9019   1.0000
   2.500   0.4253   0.02161   0.01171  -0.0459   0.8779   1.0000
   2.750   0.4567   0.02177   0.01202  -0.0454   0.8535   1.0000
   3.000   0.4868   0.02197   0.01243  -0.0446   0.8342   1.0000
   3.250   0.5112   0.02222   0.01284  -0.0429   0.8117   1.0000
   3.500   0.5354   0.02223   0.01301  -0.0401   0.7900   1.0000
   3.750   0.5567   0.02236   0.01331  -0.0373   0.7647   1.0000
   4.000   0.5766   0.02227   0.01339  -0.0335   0.7369   1.0000
   4.250   0.5937   0.02184   0.01312  -0.0281   0.7040   1.0000
   4.500   0.6113   0.02144   0.01285  -0.0233   0.6642   1.0000
   4.750   0.6244   0.02037   0.01168  -0.0164   0.5942   1.0000
   5.000   0.6421   0.01976   0.01099  -0.0118   0.4750   1.0000
   5.250   0.6523   0.02354   0.01235  -0.0097   0.1644   1.0000
   5.500   0.6748   0.02565   0.01430  -0.0090   0.1377   1.0000
   5.750   0.6992   0.02741   0.01606  -0.0078   0.1262   1.0000
   6.000   0.7251   0.02939   0.01790  -0.0065   0.1193   1.0000
   6.250   0.7541   0.03146   0.02024  -0.0055   0.1151   1.0000
   6.500   0.7822   0.03386   0.02285  -0.0048   0.1104   1.0000
   6.750   0.8084   0.03680   0.02592  -0.0045   0.1050   1.0000
   7.000   0.8349   0.04017   0.02989  -0.0041   0.1044   1.0000
   7.250   0.8593   0.04418   0.03440  -0.0038   0.1050   1.0000
   7.500   0.8802   0.04842   0.03923  -0.0037   0.1048   1.0000
   7.750   0.8963   0.05295   0.04441  -0.0038   0.1043   1.0000
   8.000   0.9128   0.05828   0.05002  -0.0039   0.1058   1.0000
   8.250   0.9076   0.06506   0.05798  -0.0060   0.1148   1.0000
   8.500   0.9084   0.07162   0.06494  -0.0079   0.1227   1.0000
   8.750   0.9073   0.07932   0.07288  -0.0104   0.1347   1.0000
   9.000   0.8596   0.08880   0.08259  -0.0206   0.1420   1.0000
   9.250   0.7798   0.11809   0.11158  -0.0590   0.3184   1.0000
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