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NASA/AMES 63A108 MOD C AIRFOIL (a63a108c-il) Xfoil prediction polar at RE=200,000 Ncrit=9


Details Polar file
Airfoil: NASA/AMES 63A108 MOD C AIRFOIL (a63a108c-il)
Reynolds number: 200,000
Max Cl/Cd: 37.14 at α=5.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-a63a108c-il-200000.txt
Download as CSV file: xf-a63a108c-il-200000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NASA/AMES 63A108 MOD C AIRFOIL                  
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.200 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.500  -0.5660   0.10690   0.10354   0.0183   1.0000   0.0501
  -9.250  -0.5895   0.09912   0.09580   0.0116   1.0000   0.0511
  -9.000  -0.7018   0.10422   0.10074   0.0263   1.0000   0.0469
  -8.750  -0.6978   0.10056   0.09710   0.0253   1.0000   0.0478
  -8.500  -0.6982   0.09586   0.09242   0.0223   1.0000   0.0490
  -8.250  -0.7060   0.08914   0.08575   0.0154   1.0000   0.0503
  -8.000  -0.7386   0.07628   0.07257  -0.0020   1.0000   0.0512
  -7.750  -0.7260   0.07298   0.06937  -0.0010   1.0000   0.0518
  -7.500  -0.7146   0.07029   0.06670  -0.0006   1.0000   0.0527
  -7.250  -0.7073   0.06591   0.06222  -0.0025   1.0000   0.0546
  -7.000  -0.7058   0.05875   0.05465  -0.0063   1.0000   0.0578
  -6.750  -0.6898   0.05595   0.05186  -0.0063   1.0000   0.0594
  -6.500  -0.6787   0.05064   0.04609  -0.0080   1.0000   0.0643
  -6.250  -0.6595   0.04796   0.04340  -0.0081   1.0000   0.0665
  -6.000  -0.6421   0.04398   0.03907  -0.0087   1.0000   0.0722
  -5.750  -0.6220   0.04112   0.03595  -0.0090   1.0000   0.0799
  -5.500  -0.5963   0.02834   0.02142  -0.0071   1.0000   0.0421
  -5.250  -0.5695   0.02504   0.01768  -0.0066   1.0000   0.0394
  -5.000  -0.5420   0.02244   0.01465  -0.0062   1.0000   0.0387
  -4.750  -0.5140   0.02066   0.01259  -0.0059   1.0000   0.0397
  -4.500  -0.4860   0.01925   0.01094  -0.0057   1.0000   0.0420
  -4.250  -0.4581   0.01805   0.00973  -0.0056   1.0000   0.0440
  -4.000  -0.4296   0.01692   0.00843  -0.0053   1.0000   0.0459
  -3.750  -0.4023   0.01582   0.00744  -0.0053   1.0000   0.0492
  -3.500  -0.3747   0.01494   0.00661  -0.0052   1.0000   0.0540
  -3.250  -0.3471   0.01404   0.00578  -0.0052   1.0000   0.0593
  -3.000  -0.3193   0.01330   0.00511  -0.0052   1.0000   0.0684
  -2.750  -0.2915   0.01252   0.00445  -0.0053   1.0000   0.0831
  -2.500  -0.2641   0.01147   0.00376  -0.0055   1.0000   0.1364
  -2.250  -0.2434   0.00916   0.00334  -0.0054   1.0000   0.5370
  -2.000  -0.2211   0.00859   0.00337  -0.0038   1.0000   0.6910
  -1.750  -0.1984   0.00834   0.00337  -0.0022   1.0000   0.7693
  -1.500  -0.1774   0.00816   0.00337  -0.0001   1.0000   0.8287
  -1.250  -0.1577   0.00807   0.00342   0.0024   1.0000   0.8852
  -1.000  -0.1271   0.00809   0.00349   0.0027   1.0000   0.9452
  -0.750  -0.0704   0.00817   0.00352  -0.0028   1.0000   0.9894
  -0.500  -0.0384   0.00816   0.00345  -0.0045   1.0000   1.0000
  -0.250  -0.0252   0.00822   0.00344  -0.0025   1.0000   1.0000
   0.000  -0.0029   0.00833   0.00349  -0.0020   0.9989   1.0000
   0.250   0.0507   0.00827   0.00340  -0.0074   0.9886   1.0000
   0.500   0.1004   0.00822   0.00333  -0.0118   0.9750   1.0000
   0.750   0.1392   0.00823   0.00333  -0.0138   0.9562   1.0000
   1.000   0.1662   0.00830   0.00337  -0.0130   0.9345   1.0000
   1.250   0.1868   0.00837   0.00341  -0.0107   0.9110   1.0000
   1.500   0.2067   0.00842   0.00342  -0.0083   0.8872   1.0000
   1.750   0.2277   0.00844   0.00340  -0.0061   0.8606   1.0000
   2.000   0.2499   0.00844   0.00336  -0.0042   0.8297   1.0000
   2.250   0.2724   0.00846   0.00329  -0.0023   0.7916   1.0000
   2.500   0.2956   0.00853   0.00321  -0.0005   0.7379   1.0000
   2.750   0.3190   0.00877   0.00313   0.0011   0.6546   1.0000
   3.000   0.3430   0.00938   0.00313   0.0023   0.5176   1.0000
   3.250   0.3688   0.01032   0.00335   0.0023   0.3673   1.0000
   3.500   0.3958   0.01115   0.00367   0.0021   0.2667   1.0000
   3.750   0.4232   0.01187   0.00404   0.0018   0.2074   1.0000
   4.000   0.4507   0.01250   0.00445   0.0016   0.1688   1.0000
   4.250   0.4781   0.01316   0.00493   0.0015   0.1420   1.0000
   4.500   0.5057   0.01374   0.00543   0.0015   0.1222   1.0000
   4.750   0.5329   0.01438   0.00596   0.0014   0.1074   1.0000
   5.000   0.5599   0.01512   0.00659   0.0015   0.0957   1.0000
   5.250   0.5868   0.01586   0.00726   0.0015   0.0862   1.0000
   5.500   0.6138   0.01654   0.00791   0.0016   0.0783   1.0000
   5.750   0.6407   0.01725   0.00862   0.0017   0.0715   1.0000
   6.000   0.6675   0.01807   0.00949   0.0019   0.0657   1.0000
   6.250   0.6941   0.01903   0.01045   0.0020   0.0610   1.0000
   6.500   0.7202   0.02013   0.01146   0.0022   0.0574   1.0000
   6.750   0.7469   0.02105   0.01257   0.0024   0.0536   1.0000
   7.000   0.7729   0.02229   0.01390   0.0026   0.0507   1.0000
   7.250   0.7990   0.02361   0.01537   0.0029   0.0484   1.0000
   7.500   0.8243   0.02521   0.01698   0.0030   0.0468   1.0000
   7.750   0.8486   0.02712   0.01942   0.0035   0.0447   1.0000
   8.000   0.8734   0.02832   0.02064   0.0036   0.0429   1.0000
   8.250   0.8956   0.03080   0.02344   0.0039   0.0419   1.0000
   8.500   0.9137   0.03443   0.02773   0.0046   0.0414   1.0000
   8.750   0.9272   0.03900   0.03293   0.0052   0.0414   1.0000
   9.000   0.9362   0.04403   0.03849   0.0058   0.0417   1.0000
   9.250   0.9411   0.04923   0.04413   0.0062   0.0419   1.0000
   9.500   0.9422   0.05462   0.04990   0.0063   0.0425   1.0000
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