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NACA 5-H-10 AIRFOIL (n5h10-il) Xfoil prediction polar at RE=200,000 Ncrit=9


Details Polar file
Airfoil: NACA 5-H-10 AIRFOIL (n5h10-il)
Reynolds number: 200,000
Max Cl/Cd: 65.83 at α=5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-n5h10-il-200000.txt
Download as CSV file: xf-n5h10-il-200000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NACA 5-H-10 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  -0.4035   0.09118   0.08833  -0.0160   0.9627   0.0419
  -9.250  -0.4005   0.08601   0.08313  -0.0198   0.9437   0.0431
  -9.000  -0.4035   0.08099   0.07808  -0.0235   0.9313   0.0445
  -8.750  -0.4144   0.07591   0.07300  -0.0282   0.9207   0.0455
  -8.500  -0.4278   0.07108   0.06812  -0.0323   0.9111   0.0457
  -8.250  -0.4432   0.06714   0.06412  -0.0339   0.9011   0.0459
  -8.000  -0.4600   0.06377   0.06061  -0.0335   0.8918   0.0462
  -7.500  -0.4747   0.05663   0.05306  -0.0319   0.8762   0.0466
  -7.250  -0.4819   0.04904   0.04540  -0.0313   0.8709   0.0475
  -7.000  -0.4710   0.04500   0.04143  -0.0309   0.8646   0.0488
  -6.750  -0.4632   0.04198   0.03834  -0.0298   0.8583   0.0504
  -6.500  -0.4550   0.03870   0.03494  -0.0289   0.8522   0.0529
  -6.250  -0.4481   0.03758   0.03310  -0.0260   0.8444   0.0595
  -6.000  -0.4599   0.04443   0.03959  -0.0275   0.8537   0.0619
  -5.750  -0.4437   0.04215   0.03731  -0.0267   0.8471   0.0648
  -5.500  -0.4263   0.04382   0.03819  -0.0233   0.8410   0.0734
  -5.250  -0.4146   0.03713   0.03164  -0.0234   0.8348   0.0767
  -5.000  -0.3965   0.03519   0.02959  -0.0221   0.8295   0.0815
  -4.750  -0.3792   0.03299   0.02704  -0.0207   0.8235   0.0916
  -4.500  -0.3609   0.03137   0.02521  -0.0195   0.8180   0.1055
  -4.250  -0.3298   0.02503   0.01770  -0.0155   0.8145   0.0540
  -4.000  -0.3029   0.02262   0.01487  -0.0146   0.8086   0.0518
  -3.750  -0.2772   0.02154   0.01358  -0.0139   0.8036   0.0550
  -3.500  -0.2505   0.02034   0.01213  -0.0131   0.7995   0.0566
  -3.250  -0.2221   0.01934   0.01095  -0.0130   0.7937   0.0582
  -3.000  -0.1951   0.01816   0.00967  -0.0127   0.7895   0.0619
  -2.750  -0.1694   0.01733   0.00883  -0.0121   0.7862   0.0648
  -2.500  -0.1427   0.01671   0.00823  -0.0120   0.7804   0.0679
  -2.250  -0.1174   0.01631   0.00777  -0.0114   0.7759   0.0727
  -2.000  -0.0949   0.01560   0.00708  -0.0102   0.7726   0.0776
  -1.750  -0.0704   0.01524   0.00674  -0.0097   0.7674   0.0846
  -1.500  -0.0468   0.01481   0.00634  -0.0089   0.7631   0.0995
  -1.250  -0.0472   0.01248   0.00592  -0.0041   0.7596   0.5442
  -1.000   0.2097   0.01343   0.00749  -0.0445   0.7619   1.0000
  -0.750   0.2329   0.01341   0.00738  -0.0437   0.7585   1.0000
  -0.500   0.2576   0.01344   0.00738  -0.0435   0.7533   1.0000
  -0.250   0.2815   0.01343   0.00731  -0.0430   0.7484   1.0000
   0.000   0.3047   0.01339   0.00721  -0.0421   0.7447   1.0000
   0.250   0.3292   0.01343   0.00724  -0.0418   0.7398   1.0000
   0.500   0.3535   0.01345   0.00724  -0.0414   0.7348   1.0000
   0.750   0.3769   0.01341   0.00717  -0.0406   0.7310   1.0000
   1.000   0.4011   0.01345   0.00721  -0.0401   0.7267   1.0000
   1.250   0.4258   0.01353   0.00732  -0.0399   0.7218   1.0000
   1.500   0.4497   0.01353   0.00733  -0.0392   0.7180   1.0000
   1.750   0.4732   0.01351   0.00729  -0.0384   0.7150   1.0000
   2.000   0.4978   0.01360   0.00746  -0.0381   0.7084   1.0000
   2.250   0.5211   0.01346   0.00731  -0.0371   0.7029   1.0000
   2.500   0.5435   0.01319   0.00706  -0.0358   0.6924   1.0000
   2.750   0.5655   0.01277   0.00662  -0.0341   0.6800   1.0000
   3.000   0.5879   0.01245   0.00627  -0.0326   0.6682   1.0000
   3.250   0.6111   0.01227   0.00608  -0.0315   0.6591   1.0000
   3.500   0.6342   0.01212   0.00597  -0.0304   0.6467   1.0000
   3.750   0.6575   0.01202   0.00595  -0.0294   0.6350   1.0000
   4.000   0.6806   0.01191   0.00587  -0.0284   0.6212   1.0000
   4.250   0.7033   0.01177   0.00576  -0.0272   0.6034   1.0000
   4.500   0.7261   0.01170   0.00576  -0.0261   0.5831   1.0000
   4.750   0.7484   0.01164   0.00574  -0.0249   0.5530   1.0000
   5.000   0.7695   0.01169   0.00570  -0.0235   0.4899   1.0000
   5.250   0.7791   0.01338   0.00624  -0.0212   0.2786   1.0000
   5.500   0.7863   0.01580   0.00759  -0.0194   0.0964   1.0000
   5.750   0.8026   0.01677   0.00846  -0.0179   0.0717   1.0000
   6.000   0.8190   0.01755   0.00926  -0.0163   0.0622   1.0000
   6.250   0.8339   0.01837   0.01008  -0.0145   0.0560   1.0000
   6.500   0.8465   0.01932   0.01110  -0.0123   0.0526   1.0000
   6.750   0.8594   0.02014   0.01198  -0.0100   0.0494   1.0000
   7.000   0.8697   0.02107   0.01290  -0.0075   0.0457   1.0000
   7.250   0.8768   0.02261   0.01444  -0.0043   0.0432   1.0000
   7.500   0.8897   0.02360   0.01554  -0.0017   0.0420   1.0000
   7.750   0.9033   0.02469   0.01673   0.0007   0.0402   1.0000
   8.000   0.9158   0.02571   0.01782   0.0032   0.0375   1.0000
   8.250   0.9297   0.02700   0.01916   0.0054   0.0359   1.0000
   8.500   0.9477   0.02899   0.02124   0.0071   0.0345   1.0000
   8.750   0.9664   0.03167   0.02411   0.0086   0.0338   1.0000
   9.000   0.9803   0.03446   0.02718   0.0107   0.0336   1.0000
   9.250   0.9839   0.03765   0.03074   0.0139   0.0325   1.0000
   9.500   0.9900   0.03950   0.03289   0.0168   0.0321   1.0000
   9.750   0.9939   0.04238   0.03609   0.0196   0.0320   1.0000
  10.000   1.0077   0.04578   0.03966   0.0212   0.0336   1.0000
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