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AH21 9% version (Andrew Hollom) (ah21-9-il) Xfoil prediction polar at RE=50,000 Ncrit=5


Details Polar file
Airfoil: AH21 9% version (Andrew Hollom) (ah21-9-il)
Reynolds number: 50,000
Max Cl/Cd: 34.53 at α=4.5°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-ah21-9-il-50000-n5.txt
Download as CSV file: xf-ah21-9-il-50000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: AH21 9% version (Andrew Hollom)                 
  
 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
  ------ -------- --------- --------- -------- -------- --------
  -8.750  -0.5000   0.09869   0.09148  -0.0266   1.0000   0.0625
  -8.500  -0.5011   0.09478   0.08765  -0.0272   1.0000   0.0584
  -8.250  -0.5185   0.08992   0.08298  -0.0321   1.0000   0.0518
  -8.000  -0.5174   0.08642   0.07953  -0.0314   1.0000   0.0506
  -7.750  -0.5193   0.08239   0.07557  -0.0328   1.0000   0.0499
  -7.500  -0.5216   0.07825   0.07145  -0.0343   1.0000   0.0495
  -7.250  -0.5228   0.07410   0.06730  -0.0355   1.0000   0.0491
  -7.000  -0.5232   0.06980   0.06295  -0.0368   1.0000   0.0492
  -6.750  -0.5216   0.06553   0.05858  -0.0377   1.0000   0.0492
  -6.500  -0.5176   0.06125   0.05407  -0.0385   1.0000   0.0493
  -6.250  -0.5110   0.05701   0.04961  -0.0389   1.0000   0.0493
  -6.000  -0.5018   0.05283   0.04515  -0.0391   1.0000   0.0488
  -5.750  -0.4898   0.04871   0.04069  -0.0391   1.0000   0.0481
  -5.500  -0.4751   0.04477   0.03634  -0.0389   1.0000   0.0476
  -5.250  -0.4579   0.04111   0.03219  -0.0386   1.0000   0.0474
  -5.000  -0.4384   0.03778   0.02835  -0.0381   1.0000   0.0476
  -4.750  -0.4171   0.03482   0.02486  -0.0374   1.0000   0.0482
  -4.500  -0.3942   0.03248   0.02189  -0.0367   1.0000   0.0514
  -4.250  -0.3718   0.03026   0.01935  -0.0360   1.0000   0.0549
  -4.000  -0.3483   0.02835   0.01714  -0.0351   1.0000   0.0567
  -3.750  -0.3243   0.02667   0.01519  -0.0341   1.0000   0.0586
  -3.500  -0.3005   0.02525   0.01346  -0.0329   1.0000   0.0611
  -3.250  -0.2768   0.02405   0.01200  -0.0316   1.0000   0.0642
  -3.000  -0.2534   0.02294   0.01076  -0.0306   1.0000   0.0684
  -2.750  -0.2299   0.02210   0.00975  -0.0297   1.0000   0.0793
  -2.500  -0.2059   0.02112   0.00883  -0.0292   1.0000   0.0993
  -2.250  -0.1807   0.01961   0.00781  -0.0292   1.0000   0.1777
  -2.000  -0.1722   0.01703   0.00782  -0.0247   1.0000   0.6967
  -1.750  -0.1301   0.01650   0.00737  -0.0256   1.0000   1.0000
  -1.500  -0.1086   0.01651   0.00700  -0.0248   1.0000   1.0000
  -1.250  -0.0868   0.01655   0.00671  -0.0240   1.0000   1.0000
  -1.000  -0.0649   0.01662   0.00650  -0.0233   1.0000   1.0000
  -0.750  -0.0428   0.01671   0.00632  -0.0226   1.0000   1.0000
  -0.500  -0.0206   0.01683   0.00623  -0.0219   1.0000   1.0000
  -0.250   0.0016   0.01697   0.00619  -0.0212   1.0000   1.0000
   0.000   0.0237   0.01713   0.00621  -0.0206   1.0000   1.0000
   0.250   0.0457   0.01732   0.00625  -0.0200   1.0000   1.0000
   0.500   0.0676   0.01753   0.00637  -0.0194   1.0000   1.0000
   0.750   0.0894   0.01777   0.00653  -0.0188   1.0000   1.0000
   1.000   0.1110   0.01804   0.00674  -0.0182   1.0000   1.0000
   1.250   0.1324   0.01833   0.00700  -0.0176   1.0000   1.0000
   1.500   0.1536   0.01866   0.00732  -0.0171   1.0000   1.0000
   1.750   0.1745   0.01901   0.00768  -0.0165   1.0000   1.0000
   2.000   0.2158   0.01961   0.00834  -0.0201   0.9891   1.0000
   2.250   0.2579   0.02018   0.00902  -0.0236   0.9774   1.0000
   2.500   0.3002   0.02067   0.00965  -0.0271   0.9643   1.0000
   2.750   0.3431   0.02106   0.01020  -0.0304   0.9491   1.0000
   3.000   0.3884   0.02130   0.01070  -0.0340   0.9317   1.0000
   3.250   0.4317   0.02132   0.01097  -0.0367   0.9098   1.0000
   3.500   0.4774   0.02106   0.01102  -0.0393   0.8832   1.0000
   3.750   0.5246   0.02040   0.01077  -0.0414   0.8492   1.0000
   4.000   0.5625   0.01962   0.01031  -0.0411   0.8018   1.0000
   4.250   0.6067   0.01877   0.00969  -0.0413   0.7160   1.0000
   4.500   0.6572   0.01903   0.00901  -0.0418   0.4659   1.0000
   4.750   0.6684   0.02085   0.00973  -0.0388   0.2933   1.0000
   5.000   0.6843   0.02264   0.01086  -0.0372   0.1956   1.0000
   5.250   0.7043   0.02427   0.01212  -0.0363   0.1440   1.0000
   5.500   0.7274   0.02579   0.01357  -0.0355   0.1180   1.0000
   5.750   0.7524   0.02720   0.01503  -0.0351   0.0972   1.0000
   6.000   0.7798   0.02891   0.01682  -0.0348   0.0864   1.0000
   6.250   0.8056   0.03060   0.01860  -0.0345   0.0760   1.0000
   6.500   0.8328   0.03236   0.02072  -0.0340   0.0677   1.0000
   6.750   0.8594   0.03479   0.02318  -0.0339   0.0635   1.0000
   7.000   0.8853   0.03754   0.02651  -0.0330   0.0606   1.0000
   7.250   0.9059   0.04018   0.02965  -0.0318   0.0565   1.0000
   7.500   0.9234   0.04229   0.03188  -0.0308   0.0518   1.0000
   7.750   0.9381   0.04571   0.03579  -0.0291   0.0498   1.0000
   8.000   0.9485   0.04953   0.04026  -0.0269   0.0488   1.0000
   8.250   0.9544   0.05360   0.04489  -0.0246   0.0480   1.0000
   8.500   0.9561   0.05781   0.04960  -0.0223   0.0476   1.0000
   8.750   0.9538   0.06215   0.05437  -0.0201   0.0473   1.0000
   9.000   0.9472   0.06654   0.05911  -0.0180   0.0471   1.0000
   9.250   0.9368   0.07089   0.06374  -0.0163   0.0469   1.0000
   9.500   0.9213   0.07515   0.06821  -0.0147   0.0469   1.0000
   9.750   0.9032   0.07946   0.07266  -0.0136   0.0471   1.0000
  10.000   0.8845   0.08433   0.07762  -0.0142   0.0475   1.0000
  10.250   0.8669   0.08995   0.08330  -0.0165   0.0480   1.0000
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