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Falcon (falcon-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: Falcon (falcon-il)
Reynolds number: 50,000
Max Cl/Cd: 22.21 at α=6°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-falcon-il-50000.txt
Download as CSV file: xf-falcon-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: Falcon                                          
  
 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
  ------ -------- --------- --------- -------- -------- --------
 -10.000  -0.3228   0.11066   0.10385  -0.0089   1.0000   0.3441
  -9.750  -0.3319   0.10849   0.10175  -0.0082   1.0000   0.3582
  -9.500  -0.3315   0.10572   0.09900  -0.0074   1.0000   0.3734
  -9.250  -0.3050   0.10078   0.09403  -0.0073   1.0000   0.3850
  -9.000  -0.3023   0.09722   0.09051  -0.0069   1.0000   0.3954
  -8.750  -0.3145   0.09450   0.08786  -0.0060   1.0000   0.4073
  -8.500  -0.6387   0.07525   0.06834  -0.0271   1.0000   0.1832
  -8.250  -0.6490   0.06867   0.06160  -0.0269   1.0000   0.1687
  -8.000  -0.6555   0.06396   0.05669  -0.0255   1.0000   0.1617
  -7.750  -0.6636   0.05933   0.05184  -0.0233   1.0000   0.1568
  -7.500  -0.6847   0.05513   0.04683  -0.0191   1.0000   0.1495
  -7.250  -0.6781   0.05142   0.04289  -0.0168   1.0000   0.1485
  -7.000  -0.6732   0.04817   0.03928  -0.0140   1.0000   0.1485
  -6.750  -0.6694   0.04556   0.03613  -0.0106   1.0000   0.1500
  -6.500  -0.6530   0.04250   0.03319  -0.0093   1.0000   0.1549
  -6.250  -0.6406   0.04006   0.03047  -0.0069   1.0000   0.1582
  -6.000  -0.6283   0.03786   0.02787  -0.0043   1.0000   0.1629
  -5.750  -0.6136   0.03564   0.02539  -0.0021   1.0000   0.1696
  -5.500  -0.5955   0.03372   0.02332  -0.0004   1.0000   0.1776
  -5.250  -0.5771   0.03182   0.02126   0.0013   1.0000   0.1881
  -5.000  -0.5570   0.03012   0.01937   0.0028   1.0000   0.2010
  -4.750  -0.5345   0.02845   0.01776   0.0039   1.0000   0.2184
  -4.500  -0.5107   0.02692   0.01638   0.0047   1.0000   0.2407
  -4.250  -0.4870   0.02555   0.01524   0.0056   1.0000   0.2693
  -4.000  -0.4659   0.02440   0.01433   0.0068   1.0000   0.3090
  -3.750  -0.4484   0.02315   0.01369   0.0085   1.0000   0.3726
  -3.500  -0.4430   0.02151   0.01355   0.0135   1.0000   0.5488
  -3.250  -0.4309   0.02200   0.01498   0.0221   1.0000   0.7805
  -3.000  -0.3616   0.02428   0.01690   0.0214   1.0000   0.8835
  -2.750  -0.1167   0.02669   0.01810  -0.0112   1.0000   0.9613
  -2.500   0.0169   0.02571   0.01662  -0.0319   1.0000   1.0000
  -2.250   0.0090   0.02548   0.01643  -0.0287   1.0000   1.0000
  -2.000  -0.0099   0.02552   0.01652  -0.0240   1.0000   1.0000
  -1.750  -0.0372   0.02576   0.01679  -0.0184   1.0000   1.0000
  -1.500  -0.0653   0.02605   0.01709  -0.0129   1.0000   1.0000
  -1.250  -0.0894   0.02632   0.01734  -0.0079   1.0000   1.0000
  -1.000  -0.0282   0.02662   0.01746  -0.0169   0.9826   1.0000
  -0.750   0.0439   0.02676   0.01740  -0.0271   0.9613   1.0000
  -0.500   0.1195   0.02675   0.01725  -0.0374   0.9417   1.0000
  -0.250   0.1849   0.02668   0.01707  -0.0454   0.9201   1.0000
   0.000   0.2530   0.02642   0.01673  -0.0534   0.8998   1.0000
   0.250   0.3118   0.02610   0.01634  -0.0591   0.8796   1.0000
   0.500   0.3432   0.02617   0.01636  -0.0601   0.8555   1.0000
   0.750   0.3797   0.02606   0.01620  -0.0615   0.8344   1.0000
   1.000   0.4129   0.02595   0.01602  -0.0619   0.8149   1.0000
   1.250   0.4327   0.02621   0.01624  -0.0605   0.7937   1.0000
   1.500   0.4541   0.02642   0.01641  -0.0591   0.7740   1.0000
   1.750   0.4758   0.02663   0.01657  -0.0577   0.7554   1.0000
   2.000   0.4969   0.02687   0.01677  -0.0561   0.7377   1.0000
   2.250   0.5172   0.02716   0.01701  -0.0544   0.7204   1.0000
   2.500   0.5370   0.02750   0.01732  -0.0526   0.7035   1.0000
   2.750   0.5563   0.02790   0.01767  -0.0508   0.6869   1.0000
   3.000   0.5752   0.02834   0.01809  -0.0489   0.6705   1.0000
   3.250   0.5936   0.02884   0.01857  -0.0469   0.6544   1.0000
   3.500   0.6119   0.02938   0.01908  -0.0450   0.6384   1.0000
   3.750   0.6299   0.02996   0.01966  -0.0431   0.6226   1.0000
   4.000   0.6477   0.03059   0.02029  -0.0411   0.6070   1.0000
   4.250   0.6653   0.03127   0.02097  -0.0392   0.5915   1.0000
   4.500   0.6830   0.03197   0.02168  -0.0372   0.5761   1.0000
   4.750   0.7007   0.03271   0.02242  -0.0353   0.5611   1.0000
   5.000   0.7189   0.03342   0.02316  -0.0335   0.5461   1.0000
   5.250   0.7377   0.03414   0.02390  -0.0317   0.5316   1.0000
   5.500   0.7579   0.03477   0.02452  -0.0299   0.5173   1.0000
   5.750   0.7806   0.03525   0.02500  -0.0284   0.5034   1.0000
   6.000   0.7998   0.03601   0.02579  -0.0267   0.4893   1.0000
   6.250   0.8127   0.03723   0.02710  -0.0246   0.4748   1.0000
   6.500   0.8234   0.03869   0.02866  -0.0224   0.4611   1.0000
   6.750   0.8358   0.04010   0.03015  -0.0204   0.4481   1.0000
   7.000   0.8545   0.04107   0.03118  -0.0187   0.4359   1.0000
   7.250   0.8837   0.04132   0.03139  -0.0178   0.4245   1.0000
   7.500   0.8750   0.04449   0.03480  -0.0147   0.4129   1.0000
   7.750   0.8815   0.04662   0.03703  -0.0126   0.4026   1.0000
   8.000   0.9209   0.04614   0.03653  -0.0123   0.3924   1.0000
   8.250   0.8716   0.05295   0.04361  -0.0080   0.3841   1.0000
   8.500   0.9207   0.05188   0.04255  -0.0080   0.3758   1.0000
   8.750   0.7909   0.06660   0.05726  -0.0041   0.3728   1.0000
   9.000   0.7061   0.08019   0.07066  -0.0080   0.3721   1.0000
   9.250   0.6790   0.08710   0.07753  -0.0098   0.3708   1.0000
   9.500   0.6634   0.09285   0.08325  -0.0112   0.3707   1.0000
   9.750   0.6579   0.09805   0.08848  -0.0126   0.3733   1.0000
  10.000   0.5774   0.11602   0.10641  -0.0255   0.4889   1.0000
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