Airfoil Tools
Search 1636 airfoils Google+
You have 0 airfoils loaded.
Your Reynold number range is 50,000 to 1,000,000. (set)

AH 79-100 A AIRFOIL (ah79100a-il) Xfoil prediction polar at RE=200,000 Ncrit=5


Details Polar file
Airfoil: AH 79-100 A AIRFOIL (ah79100a-il)
Reynolds number: 200,000
Max Cl/Cd: 81.35 at α=4.75°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-ah79100a-il-200000-n5.txt
Download as CSV file: xf-ah79100a-il-200000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: AH 79-100 A 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 =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.250  -0.3934   0.08694   0.08335  -0.0425   1.0000   0.0175
  -9.000  -0.4020   0.08240   0.07887  -0.0437   1.0000   0.0173
  -8.750  -0.4157   0.07770   0.07427  -0.0444   1.0000   0.0171
  -8.500  -0.5117   0.04916   0.04567  -0.0674   0.9924   0.0155
  -8.250  -0.4928   0.03664   0.03223  -0.0877   0.9815   0.0155
  -8.000  -0.4686   0.03158   0.02665  -0.0935   0.9758   0.0159
  -7.750  -0.4399   0.02872   0.02345  -0.0969   0.9714   0.0164
  -7.500  -0.4078   0.02629   0.02066  -0.1001   0.9685   0.0170
  -7.250  -0.3795   0.02421   0.01821  -0.1018   0.9636   0.0179
  -7.000  -0.3487   0.02225   0.01589  -0.1035   0.9596   0.0190
  -6.750  -0.3153   0.02073   0.01401  -0.1054   0.9568   0.0206
  -6.500  -0.2817   0.01963   0.01285  -0.1074   0.9544   0.0225
  -6.250  -0.2541   0.01878   0.01185  -0.1078   0.9487   0.0243
  -6.000  -0.2226   0.01775   0.01062  -0.1089   0.9445   0.0262
  -5.750  -0.1894   0.01683   0.00964  -0.1104   0.9413   0.0288
  -5.500  -0.1594   0.01620   0.00894  -0.1111   0.9361   0.0316
  -5.250  -0.1290   0.01557   0.00816  -0.1117   0.9304   0.0343
  -5.000  -0.0962   0.01477   0.00733  -0.1130   0.9261   0.0380
  -4.750  -0.0676   0.01424   0.00673  -0.1133   0.9194   0.0415
  -4.500  -0.0371   0.01371   0.00612  -0.1139   0.9138   0.0466
  -4.250  -0.0053   0.01322   0.00558  -0.1148   0.9096   0.0544
  -4.000   0.0218   0.01281   0.00515  -0.1147   0.9024   0.0643
  -3.750   0.0521   0.01240   0.00474  -0.1152   0.8973   0.0814
  -3.500   0.0807   0.01198   0.00440  -0.1155   0.8912   0.1128
  -3.250   0.1088   0.01149   0.00412  -0.1158   0.8846   0.1709
  -3.000   0.1380   0.01101   0.00390  -0.1163   0.8789   0.2522
  -2.750   0.1650   0.01077   0.00379  -0.1162   0.8710   0.3084
  -2.250   0.2212   0.01049   0.00364  -0.1160   0.8570   0.3955
  -2.000   0.2502   0.01040   0.00355  -0.1160   0.8508   0.4253
  -1.750   0.2770   0.01035   0.00352  -0.1156   0.8423   0.4497
  -1.500   0.3056   0.01029   0.00345  -0.1155   0.8354   0.4748
  -1.250   0.3324   0.01025   0.00342  -0.1151   0.8257   0.4927
  -1.000   0.3602   0.01021   0.00335  -0.1149   0.8173   0.5094
  -0.750   0.3877   0.01018   0.00331  -0.1146   0.8082   0.5251
  -0.500   0.4150   0.01016   0.00329  -0.1143   0.7991   0.5392
  -0.250   0.4429   0.01013   0.00324  -0.1140   0.7904   0.5547
   0.000   0.4697   0.01011   0.00323  -0.1136   0.7794   0.5696
   0.250   0.4968   0.01009   0.00321  -0.1132   0.7680   0.5833
   0.500   0.5239   0.01007   0.00320  -0.1129   0.7566   0.5975
   0.750   0.5510   0.01006   0.00319  -0.1125   0.7454   0.6136
   1.000   0.5779   0.01006   0.00322  -0.1121   0.7334   0.6304
   1.250   0.6046   0.01006   0.00325  -0.1117   0.7206   0.6481
   1.750   0.6570   0.01007   0.00332  -0.1106   0.6916   0.6877
   2.000   0.6827   0.01008   0.00338  -0.1099   0.6768   0.7132
   2.250   0.7082   0.01008   0.00346  -0.1092   0.6622   0.7462
   2.500   0.7332   0.01004   0.00352  -0.1083   0.6467   0.7978
   2.750   0.7635   0.00987   0.00351  -0.1084   0.6290   1.0000
   3.000   0.7900   0.01006   0.00362  -0.1081   0.6114   1.0000
   3.250   0.8164   0.01026   0.00377  -0.1077   0.5946   1.0000
   3.500   0.8423   0.01048   0.00393  -0.1073   0.5757   1.0000
   3.750   0.8674   0.01075   0.00409  -0.1067   0.5534   1.0000
   4.000   0.8924   0.01102   0.00428  -0.1061   0.5298   1.0000
   4.250   0.9174   0.01131   0.00451  -0.1055   0.5094   1.0000
   4.500   0.9426   0.01159   0.00476  -0.1049   0.4903   1.0000
   4.750   0.9672   0.01189   0.00501  -0.1043   0.4682   1.0000
   5.000   0.9912   0.01223   0.00531  -0.1036   0.4435   1.0000
   5.250   1.0144   0.01262   0.00561  -0.1027   0.4160   1.0000
   5.500   1.0369   0.01307   0.00595  -0.1018   0.3843   1.0000
   5.750   1.0588   0.01357   0.00633  -0.1008   0.3506   1.0000
   6.000   1.0800   0.01412   0.00678  -0.0997   0.3172   1.0000
   6.250   1.1011   0.01471   0.00725  -0.0987   0.2829   1.0000
   6.500   1.1211   0.01539   0.00778  -0.0975   0.2462   1.0000
   6.750   1.1388   0.01628   0.00841  -0.0961   0.1997   1.0000
   7.000   1.1557   0.01726   0.00912  -0.0947   0.1536   1.0000
   7.250   1.1732   0.01819   0.00988  -0.0933   0.1192   1.0000
   7.500   1.1913   0.01905   0.01062  -0.0920   0.0937   1.0000
   7.750   1.2089   0.01993   0.01139  -0.0906   0.0721   1.0000
   8.000   1.2258   0.02084   0.01222  -0.0891   0.0545   1.0000
   8.250   1.2418   0.02180   0.01313  -0.0875   0.0420   1.0000
   8.500   1.2570   0.02280   0.01414  -0.0857   0.0328   1.0000
   8.750   1.2721   0.02374   0.01513  -0.0839   0.0271   1.0000
   9.000   1.2844   0.02478   0.01621  -0.0817   0.0234   1.0000
   9.250   1.2967   0.02574   0.01726  -0.0795   0.0212   1.0000
   9.500   1.3063   0.02690   0.01848  -0.0771   0.0196   1.0000
   9.750   1.3138   0.02821   0.01989  -0.0745   0.0185   1.0000
  10.000   1.3224   0.02946   0.02130  -0.0722   0.0177   1.0000
  10.250   1.3298   0.03085   0.02281  -0.0699   0.0169   1.0000
  10.500   1.3365   0.03234   0.02442  -0.0677   0.0163   1.0000
  10.750   1.3429   0.03391   0.02610  -0.0657   0.0157   1.0000
  11.000   1.3485   0.03560   0.02789  -0.0639   0.0151   1.0000
  11.250   1.3520   0.03755   0.02993  -0.0620   0.0146   1.0000
  11.500   1.3523   0.03993   0.03238  -0.0602   0.0140   1.0000
  11.750   1.3583   0.04183   0.03445  -0.0588   0.0136   1.0000
  12.000   1.3632   0.04392   0.03670  -0.0576   0.0132   1.0000
  12.250   1.3663   0.04630   0.03926  -0.0564   0.0128   1.0000
  12.500   1.3689   0.04880   0.04192  -0.0553   0.0126   1.0000
  12.750   1.3708   0.05146   0.04474  -0.0544   0.0123   1.0000
  13.000   1.3718   0.05430   0.04774  -0.0537   0.0120   1.0000
  13.250   1.3718   0.05734   0.05095  -0.0532   0.0118   1.0000
  13.500   1.3709   0.06060   0.05437  -0.0529   0.0117   1.0000
  13.750   1.3686   0.06409   0.05803  -0.0529   0.0115   1.0000
  14.000   1.3651   0.06784   0.06196  -0.0532   0.0113   1.0000
  14.250   1.3605   0.07187   0.06617  -0.0538   0.0112   1.0000
  14.500   1.3546   0.07625   0.07072  -0.0547   0.0111   1.0000
  14.750   1.3475   0.08098   0.07563  -0.0559   0.0110   1.0000
  15.000   1.3392   0.08610   0.08093  -0.0576   0.0109   1.0000
  15.250   1.3298   0.09159   0.08661  -0.0597   0.0108   1.0000
  15.500   1.3191   0.09751   0.09271  -0.0623   0.0107   1.0000
  15.750   1.3074   0.10390   0.09928  -0.0654   0.0107   1.0000
  16.000   1.2947   0.11077   0.10634  -0.0690   0.0106   1.0000
  16.250   1.2808   0.11817   0.11393  -0.0732   0.0106   1.0000
  16.500   1.2658   0.12615   0.12211  -0.0780   0.0106   1.0000
  16.750   1.2492   0.13500   0.13117  -0.0836   0.0106   1.0000
  17.000   1.2282   0.14574   0.14214  -0.0906   0.0107   1.0000
  17.250   1.1659   0.17125   0.16812  -0.1076   0.0113   1.0000
<< Back to AH 79-100 A AIRFOIL (ah79100a-il)

Polar data table (+)

Polar graphs


<< Back to AH 79-100 A AIRFOIL (ah79100a-il)