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 V23010-1.58 AIRFOIL (v23010-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: BOEING-VERTOL V23010-1.58 AIRFOIL (v23010-il)
Reynolds number: 50,000
Max Cl/Cd: 23.75 at α=4.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-v23010-il-50000.txt
Download as CSV file: xf-v23010-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: BOEING-VERTOL V23010-1.58 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
  ------ -------- --------- --------- -------- -------- --------
  -9.000  -0.4672   0.10926   0.10205   0.0134   1.0000   0.3180
  -8.750  -0.4610   0.10660   0.09943   0.0142   1.0000   0.3391
  -8.500  -0.4679   0.10499   0.09791   0.0147   1.0000   0.3590
  -8.250  -0.4388   0.10058   0.09350   0.0165   1.0000   0.3857
  -8.000  -0.4393   0.09844   0.09144   0.0178   1.0000   0.4108
  -7.750  -0.4295   0.09613   0.08918   0.0196   1.0000   0.4403
  -7.500  -0.4035   0.09215   0.08520   0.0209   1.0000   0.4700
  -7.250  -0.3864   0.08896   0.08205   0.0223   1.0000   0.5004
  -7.000  -0.3707   0.08595   0.07908   0.0237   1.0000   0.5314
  -6.750  -0.3588   0.08319   0.07637   0.0251   1.0000   0.5612
  -6.250  -0.3264   0.07700   0.07026   0.0265   1.0000   0.6121
  -6.000  -0.3123   0.07396   0.06726   0.0268   1.0000   0.6313
  -5.750  -0.3021   0.07109   0.06445   0.0271   1.0000   0.6462
  -5.000  -0.4839   0.05155   0.04507  -0.0038   1.0000   0.3239
  -4.750  -0.4679   0.04353   0.03572  -0.0120   1.0000   0.2287
  -4.500  -0.4457   0.04020   0.03167  -0.0119   1.0000   0.1965
  -4.250  -0.4241   0.03759   0.02859  -0.0108   1.0000   0.1789
  -4.000  -0.4023   0.03570   0.02611  -0.0093   1.0000   0.1651
  -3.750  -0.3816   0.03350   0.02375  -0.0080   1.0000   0.1597
  -3.500  -0.3602   0.03190   0.02176  -0.0064   1.0000   0.1543
  -3.250  -0.3385   0.03095   0.02039  -0.0047   1.0000   0.1502
  -3.000  -0.3167   0.02928   0.01864  -0.0034   1.0000   0.1478
  -2.750  -0.2946   0.02794   0.01719  -0.0021   1.0000   0.1449
  -2.500  -0.2722   0.02684   0.01595  -0.0007   1.0000   0.1425
  -2.250  -0.2499   0.02591   0.01490   0.0006   1.0000   0.1410
  -2.000  -0.2276   0.02510   0.01404   0.0020   1.0000   0.1406
  -1.750  -0.2057   0.02437   0.01333   0.0033   1.0000   0.1424
  -1.500  -0.1845   0.02380   0.01276   0.0045   1.0000   0.1460
  -1.250  -0.1643   0.02339   0.01228   0.0057   1.0000   0.1498
  -1.000  -0.1470   0.02292   0.01185   0.0072   1.0000   0.1534
  -0.750  -0.1305   0.02256   0.01153   0.0085   1.0000   0.1584
  -0.500  -0.1139   0.02238   0.01132   0.0097   1.0000   0.1656
  -0.250  -0.0975   0.02213   0.01118   0.0105   1.0000   0.1799
   0.000  -0.0273   0.01929   0.01173   0.0045   0.9992   1.0000
   0.250   0.0757   0.02014   0.01184  -0.0093   0.9648   1.0000
   0.500   0.1857   0.02025   0.01167  -0.0240   0.9262   1.0000
   0.750   0.2889   0.01954   0.01088  -0.0357   0.8782   1.0000
   1.000   0.3417   0.01908   0.01027  -0.0377   0.8188   1.0000
   1.250   0.3697   0.01895   0.00986  -0.0351   0.7609   1.0000
   1.500   0.3882   0.01927   0.00987  -0.0317   0.7051   1.0000
   1.750   0.4071   0.01967   0.00997  -0.0289   0.6598   1.0000
   2.000   0.4276   0.02012   0.01017  -0.0268   0.6233   1.0000
   2.250   0.4495   0.02061   0.01043  -0.0251   0.5947   1.0000
   2.500   0.4714   0.02119   0.01084  -0.0237   0.5701   1.0000
   2.750   0.4941   0.02182   0.01129  -0.0225   0.5504   1.0000
   3.000   0.5168   0.02249   0.01184  -0.0213   0.5327   1.0000
   3.250   0.5391   0.02321   0.01247  -0.0202   0.5162   1.0000
   3.500   0.5612   0.02399   0.01322  -0.0192   0.5010   1.0000
   3.750   0.5830   0.02480   0.01400  -0.0181   0.4861   1.0000
   4.000   0.6046   0.02564   0.01484  -0.0171   0.4718   1.0000
   4.250   0.6262   0.02651   0.01569  -0.0160   0.4580   1.0000
   4.500   0.6481   0.02742   0.01658  -0.0149   0.4453   1.0000
   4.750   0.6712   0.02826   0.01734  -0.0138   0.4331   1.0000
   5.000   0.6913   0.02943   0.01864  -0.0129   0.4205   1.0000
   5.250   0.7110   0.03074   0.02006  -0.0119   0.4087   1.0000
   5.500   0.7328   0.03194   0.02126  -0.0109   0.3982   1.0000
   5.750   0.7520   0.03345   0.02293  -0.0100   0.3878   1.0000
   6.000   0.7690   0.03528   0.02495  -0.0091   0.3778   1.0000
   6.250   0.7931   0.03640   0.02598  -0.0082   0.3683   1.0000
   6.500   0.8024   0.03906   0.02903  -0.0073   0.3587   1.0000
   6.750   0.8264   0.04047   0.03042  -0.0065   0.3517   1.0000
   7.000   0.8270   0.04435   0.03476  -0.0060   0.3452   1.0000
   7.250   0.8417   0.04663   0.03716  -0.0052   0.3379   1.0000
   7.500   0.8478   0.04998   0.04069  -0.0046   0.3321   1.0000
   7.750   0.8311   0.05596   0.04697  -0.0052   0.3294   1.0000
   8.000   0.8026   0.06339   0.05455  -0.0070   0.3301   1.0000
   8.250   0.7778   0.07064   0.06184  -0.0094   0.3325   1.0000
   8.500   0.7620   0.07693   0.06812  -0.0114   0.3354   1.0000
   8.750   0.5794   0.09999   0.09079  -0.0343   0.5033   1.0000
<< Back to BOEING-VERTOL V23010-1.58 AIRFOIL (v23010-il)

Polar data table (+)

Polar graphs


<< Back to BOEING-VERTOL V23010-1.58 AIRFOIL (v23010-il)