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NLR-1T AIRFOIL (nlr1t-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: NLR-1T AIRFOIL (nlr1t-il)
Reynolds number: 50,000
Max Cl/Cd: 29.6 at α=4°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-nlr1t-il-50000.txt
Download as CSV file: xf-nlr1t-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NLR-1T 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
  ------ -------- --------- --------- -------- -------- --------
  -8.750  -0.5108   0.10563   0.09885  -0.0017   1.0000   0.2815
  -8.500  -0.5138   0.10288   0.09618  -0.0015   1.0000   0.2979
  -8.250  -0.5226   0.10051   0.09390  -0.0014   1.0000   0.3139
  -8.000  -0.4949   0.09575   0.08909   0.0012   1.0000   0.3370
  -7.750  -0.5113   0.09410   0.08757   0.0019   1.0000   0.3578
  -7.500  -0.4946   0.09043   0.08391   0.0044   1.0000   0.3861
  -7.250  -0.4883   0.08757   0.08110   0.0069   1.0000   0.4159
  -7.000  -0.4677   0.08380   0.07733   0.0097   1.0000   0.4501
  -6.750  -0.4622   0.08150   0.07509   0.0131   1.0000   0.4893
  -6.500  -0.4482   0.07875   0.07237   0.0167   1.0000   0.5339
  -6.250  -0.4071   0.07433   0.06787   0.0185   1.0000   0.5872
  -5.000  -0.5425   0.04968   0.04221  -0.0092   1.0000   0.2079
  -4.750  -0.5202   0.04482   0.03616  -0.0089   1.0000   0.1522
  -4.500  -0.5003   0.04131   0.03217  -0.0073   1.0000   0.1360
  -4.250  -0.4798   0.03892   0.02888  -0.0050   1.0000   0.1244
  -4.000  -0.4589   0.03616   0.02589  -0.0036   1.0000   0.1229
  -3.750  -0.4365   0.03384   0.02322  -0.0022   1.0000   0.1217
  -3.500  -0.4119   0.03167   0.02072  -0.0010   1.0000   0.1198
  -3.250  -0.3857   0.02974   0.01842   0.0002   1.0000   0.1183
  -3.000  -0.3583   0.02805   0.01644   0.0011   1.0000   0.1187
  -2.750  -0.3294   0.02659   0.01474   0.0019   1.0000   0.1218
  -2.500  -0.2991   0.02516   0.01330   0.0021   1.0000   0.1314
  -2.250  -0.1320   0.01903   0.01034  -0.0191   1.0000   1.0000
  -2.000  -0.1191   0.01896   0.00982  -0.0164   1.0000   1.0000
  -1.750  -0.1056   0.01892   0.00942  -0.0140   1.0000   1.0000
  -1.500  -0.0914   0.01891   0.00912  -0.0117   1.0000   1.0000
  -1.250  -0.0764   0.01893   0.00888  -0.0096   1.0000   1.0000
  -1.000  -0.0608   0.01896   0.00869  -0.0076   1.0000   1.0000
  -0.750  -0.0447   0.01902   0.00855  -0.0057   1.0000   1.0000
  -0.500  -0.0283   0.01910   0.00844  -0.0038   1.0000   1.0000
  -0.250  -0.0116   0.01921   0.00840  -0.0021   1.0000   1.0000
   0.000   0.0055   0.01933   0.00839  -0.0004   1.0000   1.0000
   0.250   0.0230   0.01948   0.00843   0.0012   1.0000   1.0000
   0.500   0.0407   0.01965   0.00850   0.0027   1.0000   1.0000
   0.750   0.0586   0.01985   0.00863   0.0041   1.0000   1.0000
   1.000   0.0767   0.02008   0.00880   0.0054   1.0000   1.0000
   1.250   0.0949   0.02033   0.00903   0.0066   1.0000   1.0000
   1.500   0.1131   0.02062   0.00930   0.0078   1.0000   1.0000
   1.750   0.1313   0.02095   0.00963   0.0089   1.0000   1.0000
   2.000   0.1493   0.02133   0.01003   0.0099   1.0000   1.0000
   2.250   0.1670   0.02176   0.01050   0.0109   1.0000   1.0000
   2.500   0.1843   0.02226   0.01107   0.0117   1.0000   1.0000
   2.750   0.2009   0.02284   0.01174   0.0125   1.0000   1.0000
   3.000   0.2165   0.02355   0.01254   0.0132   1.0000   1.0000
   3.250   0.3358   0.02479   0.01423  -0.0054   0.9547   1.0000
   3.500   0.4659   0.02393   0.01410  -0.0225   0.8879   1.0000
   3.750   0.5419   0.02095   0.01170  -0.0249   0.7580   1.0000
   4.000   0.6014   0.02032   0.00960  -0.0239   0.5321   1.0000
   4.250   0.6265   0.02166   0.01044  -0.0228   0.4719   1.0000
   4.500   0.6544   0.02281   0.01137  -0.0225   0.4366   1.0000
   4.750   0.6812   0.02390   0.01239  -0.0221   0.4096   1.0000
   5.000   0.7077   0.02503   0.01353  -0.0216   0.3882   1.0000
   5.250   0.7322   0.02619   0.01476  -0.0209   0.3680   1.0000
   5.500   0.7564   0.02743   0.01601  -0.0201   0.3492   1.0000
   5.750   0.7779   0.02868   0.01735  -0.0189   0.3278   1.0000
   6.000   0.7988   0.02987   0.01846  -0.0176   0.3041   1.0000
   6.250   0.8176   0.03124   0.02004  -0.0160   0.2802   1.0000
   6.500   0.8371   0.03280   0.02160  -0.0145   0.2569   1.0000
   6.750   0.8553   0.03469   0.02367  -0.0128   0.2349   1.0000
   7.000   0.8747   0.03687   0.02590  -0.0113   0.2171   1.0000
   7.250   0.8922   0.03904   0.02815  -0.0097   0.2012   1.0000
   7.500   0.9059   0.04147   0.03105  -0.0077   0.1892   1.0000
   7.750   0.9163   0.04483   0.03491  -0.0055   0.1827   1.0000
   8.000   0.9248   0.04828   0.03882  -0.0033   0.1779   1.0000
   8.250   0.9416   0.05117   0.04174  -0.0020   0.1721   1.0000
   8.500   0.9363   0.05544   0.04664   0.0009   0.1695   1.0000
   8.750   0.9303   0.05982   0.05145   0.0033   0.1671   1.0000
   9.000   0.9213   0.06452   0.05648   0.0053   0.1662   1.0000
   9.250   0.9067   0.06987   0.06208   0.0068   0.1675   1.0000
   9.500   0.8915   0.07535   0.06772   0.0078   0.1697   1.0000
   9.750   0.8858   0.08062   0.07306   0.0082   0.1718   1.0000
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