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NASA/AMES 63A108 MOD C AIRFOIL (a63a108c-il) Xfoil prediction polar at RE=50,000 Ncrit=5


Details Polar file
Airfoil: NASA/AMES 63A108 MOD C AIRFOIL (a63a108c-il)
Reynolds number: 50,000
Max Cl/Cd: 22.81 at α=5.25°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-a63a108c-il-50000-n5.txt
Download as CSV file: xf-a63a108c-il-50000-n5.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 =     0.050 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.500  -0.6873   0.11414   0.10692   0.0226   1.0000   0.0565
  -9.250  -0.6845   0.10963   0.10244   0.0208   1.0000   0.0557
  -9.000  -0.6841   0.10470   0.09755   0.0182   1.0000   0.0549
  -8.500  -0.6990   0.09019   0.08309   0.0053   1.0000   0.0509
  -8.250  -0.7024   0.08494   0.07783   0.0020   1.0000   0.0505
  -8.000  -0.7046   0.07949   0.07230  -0.0012   1.0000   0.0501
  -7.750  -0.7057   0.07403   0.06671  -0.0040   1.0000   0.0497
  -7.500  -0.7043   0.06864   0.06112  -0.0064   1.0000   0.0494
  -7.250  -0.6999   0.06336   0.05556  -0.0082   1.0000   0.0490
  -7.000  -0.6918   0.05826   0.05010  -0.0095   1.0000   0.0485
  -6.750  -0.6802   0.05340   0.04480  -0.0104   1.0000   0.0482
  -6.500  -0.6651   0.04895   0.03987  -0.0108   1.0000   0.0482
  -6.250  -0.6467   0.04502   0.03544  -0.0110   1.0000   0.0486
  -6.000  -0.6261   0.04145   0.03128  -0.0111   1.0000   0.0503
  -5.750  -0.6035   0.03855   0.02795  -0.0110   1.0000   0.0523
  -5.500  -0.5792   0.03599   0.02504  -0.0108   1.0000   0.0542
  -5.250  -0.5533   0.03336   0.02193  -0.0105   1.0000   0.0558
  -5.000  -0.5273   0.03134   0.01969  -0.0102   1.0000   0.0589
  -4.750  -0.5003   0.02948   0.01747  -0.0098   1.0000   0.0630
  -4.500  -0.4738   0.02782   0.01579  -0.0094   1.0000   0.0666
  -4.250  -0.4471   0.02635   0.01421  -0.0089   1.0000   0.0723
  -4.000  -0.4205   0.02502   0.01277  -0.0083   1.0000   0.0790
  -3.750  -0.3945   0.02381   0.01154  -0.0077   1.0000   0.0879
  -3.500  -0.3687   0.02264   0.01041  -0.0073   1.0000   0.1001
  -3.250  -0.3423   0.02149   0.00927  -0.0071   1.0000   0.1198
  -3.000  -0.3174   0.02011   0.00820  -0.0069   1.0000   0.1589
  -2.750  -0.2964   0.01788   0.00723  -0.0066   1.0000   0.3413
  -2.500  -0.2835   0.01649   0.00728  -0.0023   1.0000   0.6458
  -2.250  -0.2657   0.01611   0.00727   0.0020   1.0000   0.7861
  -2.000  -0.2234   0.01604   0.00727   0.0024   1.0000   0.9158
  -1.750  -0.1333   0.01609   0.00690  -0.0084   1.0000   0.9964
  -1.500  -0.1096   0.01579   0.00643  -0.0087   1.0000   1.0000
  -1.250  -0.0906   0.01555   0.00604  -0.0079   1.0000   1.0000
  -1.000  -0.0715   0.01537   0.00575  -0.0070   1.0000   1.0000
  -0.750  -0.0526   0.01526   0.00553  -0.0060   1.0000   1.0000
  -0.500  -0.0337   0.01521   0.00538  -0.0049   1.0000   1.0000
  -0.250  -0.0149   0.01521   0.00529  -0.0038   1.0000   1.0000
   0.000   0.0040   0.01525   0.00527  -0.0026   1.0000   1.0000
   0.250   0.0230   0.01534   0.00530  -0.0014   1.0000   1.0000
   0.500   0.0425   0.01546   0.00539  -0.0004   1.0000   1.0000
   0.750   0.0624   0.01562   0.00553   0.0006   1.0000   1.0000
   1.000   0.0826   0.01581   0.00572   0.0014   1.0000   1.0000
   1.250   0.1032   0.01605   0.00598   0.0021   1.0000   1.0000
   1.500   0.1361   0.01632   0.00630   0.0004   0.9928   1.0000
   1.750   0.2046   0.01648   0.00661  -0.0076   0.9637   1.0000
   2.000   0.2584   0.01657   0.00686  -0.0121   0.9276   1.0000
   2.250   0.2970   0.01664   0.00705  -0.0131   0.8850   1.0000
   2.500   0.3247   0.01669   0.00716  -0.0117   0.8356   1.0000
   2.750   0.3477   0.01672   0.00721  -0.0092   0.7757   1.0000
   3.000   0.3684   0.01679   0.00717  -0.0061   0.6957   1.0000
   3.250   0.3872   0.01710   0.00708  -0.0025   0.5797   1.0000
   3.500   0.4061   0.01795   0.00717   0.0003   0.4433   1.0000
   3.750   0.4278   0.01909   0.00765   0.0013   0.3380   1.0000
   4.000   0.4515   0.02018   0.00832   0.0018   0.2712   1.0000
   4.250   0.4760   0.02122   0.00912   0.0021   0.2266   1.0000
   4.500   0.5016   0.02223   0.00997   0.0023   0.1939   1.0000
   4.750   0.5272   0.02326   0.01088   0.0026   0.1696   1.0000
   5.000   0.5530   0.02429   0.01190   0.0029   0.1498   1.0000
   5.250   0.5787   0.02537   0.01299   0.0033   0.1342   1.0000
   5.500   0.6045   0.02653   0.01417   0.0037   0.1214   1.0000
   5.750   0.6303   0.02780   0.01549   0.0041   0.1108   1.0000
   6.000   0.6558   0.02916   0.01695   0.0045   0.1020   1.0000
   6.250   0.6810   0.03056   0.01836   0.0048   0.0947   1.0000
   6.500   0.7063   0.03221   0.02025   0.0051   0.0879   1.0000
   6.750   0.7311   0.03416   0.02249   0.0054   0.0823   1.0000
   7.000   0.7549   0.03583   0.02421   0.0056   0.0778   1.0000
   7.250   0.7776   0.03854   0.02749   0.0059   0.0734   1.0000
   7.500   0.7997   0.04057   0.02965   0.0061   0.0700   1.0000
   7.750   0.8179   0.04415   0.03390   0.0063   0.0668   1.0000
   8.000   0.8363   0.04693   0.03697   0.0065   0.0645   1.0000
   8.250   0.8491   0.05092   0.04147   0.0065   0.0625   1.0000
   8.500   0.8556   0.05583   0.04696   0.0061   0.0606   1.0000
   8.750   0.8602   0.06041   0.05193   0.0057   0.0594   1.0000
   9.000   0.8584   0.06574   0.05762   0.0048   0.0589   1.0000
   9.250   0.8468   0.07214   0.06434   0.0028   0.0590   1.0000
   9.500   0.8244   0.07974   0.07216  -0.0010   0.0596   1.0000
   9.750   0.7980   0.08865   0.08112  -0.0075   0.0606   1.0000
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