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

A18 (original) (a18-il) Xfoil prediction polar at RE=500,000 Ncrit=9


Details Polar file
Airfoil: A18 (original) (a18-il)
Reynolds number: 500,000
Max Cl/Cd: 111.88 at α=2°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-a18-il-500000.txt
Download as CSV file: xf-a18-il-500000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: A18 (original)                                  
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.500 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -10.000  -0.4139   0.09993   0.09759  -0.0270   1.0000   0.0254
  -9.750  -0.4171   0.09608   0.09377  -0.0302   1.0000   0.0266
  -9.000  -0.5381   0.02452   0.02100  -0.0866   1.0000   0.0146
  -8.750  -0.5136   0.02169   0.01769  -0.0882   1.0000   0.0152
  -8.500  -0.4879   0.01864   0.01421  -0.0899   1.0000   0.0166
  -8.250  -0.4640   0.01789   0.01334  -0.0896   1.0000   0.0176
  -8.000  -0.4384   0.01685   0.01208  -0.0897   0.9999   0.0187
  -7.750  -0.4031   0.01579   0.01077  -0.0916   0.9987   0.0200
  -7.500  -0.3678   0.01441   0.00918  -0.0937   0.9975   0.0224
  -7.250  -0.3315   0.01419   0.00891  -0.0955   0.9958   0.0252
  -7.000  -0.2951   0.01353   0.00813  -0.0975   0.9947   0.0291
  -6.750  -0.2599   0.01357   0.00816  -0.0989   0.9925   0.0333
  -6.500  -0.2252   0.01309   0.00759  -0.1005   0.9903   0.0380
  -6.250  -0.1895   0.01325   0.00775  -0.1022   0.9879   0.0423
  -6.000  -0.1528   0.01336   0.00777  -0.1039   0.9859   0.0454
  -5.750  -0.1164   0.01235   0.00667  -0.1060   0.9847   0.0497
  -5.500  -0.0790   0.01216   0.00645  -0.1081   0.9834   0.0534
  -5.250  -0.0470   0.01196   0.00617  -0.1089   0.9794   0.0565
  -5.000  -0.0121   0.01144   0.00557  -0.1105   0.9767   0.0590
  -4.750   0.0240   0.01071   0.00481  -0.1123   0.9748   0.0623
  -4.500   0.0610   0.01035   0.00443  -0.1143   0.9731   0.0657
  -4.250   0.0985   0.00998   0.00402  -0.1163   0.9717   0.0686
  -4.000   0.1330   0.00970   0.00369  -0.1177   0.9688   0.0708
  -3.750   0.1647   0.00922   0.00322  -0.1185   0.9637   0.0757
  -3.500   0.1985   0.00893   0.00294  -0.1197   0.9600   0.0812
  -3.250   0.2328   0.00864   0.00265  -0.1209   0.9567   0.0891
  -3.000   0.2614   0.00838   0.00246  -0.1210   0.9488   0.1036
  -2.750   0.2932   0.00805   0.00228  -0.1217   0.9434   0.1440
  -2.500   0.3215   0.00783   0.00219  -0.1217   0.9344   0.1847
  -2.250   0.3509   0.00766   0.00211  -0.1219   0.9262   0.2206
  -2.000   0.3792   0.00750   0.00201  -0.1218   0.9158   0.2505
  -1.750   0.4071   0.00732   0.00195  -0.1217   0.9044   0.2893
  -1.500   0.4350   0.00718   0.00196  -0.1216   0.8926   0.3580
  -1.250   0.4626   0.00710   0.00192  -0.1213   0.8786   0.3961
  -1.000   0.4899   0.00700   0.00187  -0.1211   0.8620   0.4298
  -0.750   0.5172   0.00676   0.00190  -0.1209   0.8446   0.5328
  -0.500   0.5402   0.00602   0.00198  -0.1197   0.8276   0.8154
  -0.250   0.5633   0.00578   0.00185  -0.1181   0.8081   1.0000
   0.000   0.5907   0.00590   0.00185  -0.1179   0.7887   1.0000
   0.250   0.6182   0.00602   0.00187  -0.1176   0.7705   1.0000
   0.500   0.6457   0.00615   0.00191  -0.1174   0.7512   1.0000
   0.750   0.6730   0.00630   0.00198  -0.1172   0.7310   1.0000
   1.000   0.7001   0.00646   0.00204  -0.1170   0.7086   1.0000
   1.250   0.7273   0.00663   0.00212  -0.1168   0.6873   1.0000
   1.500   0.7545   0.00680   0.00222  -0.1166   0.6655   1.0000
   2.000   0.8078   0.00722   0.00244  -0.1160   0.6116   1.0000
   2.250   0.8339   0.00750   0.00258  -0.1156   0.5803   1.0000
   2.500   0.8598   0.00781   0.00277  -0.1153   0.5502   1.0000
   2.750   0.8855   0.00814   0.00297  -0.1149   0.5163   1.0000
   3.000   0.9113   0.00845   0.00315  -0.1145   0.4699   1.0000
   3.250   0.9346   0.00906   0.00340  -0.1138   0.4070   1.0000
   3.500   0.9591   0.00958   0.00374  -0.1134   0.3541   1.0000
   3.750   0.9791   0.01072   0.00425  -0.1124   0.2283   1.0000
   4.000   1.0028   0.01138   0.00469  -0.1119   0.1866   1.0000
   4.250   1.0274   0.01188   0.00507  -0.1114   0.1571   1.0000
   4.500   1.0518   0.01241   0.00546  -0.1110   0.1209   1.0000
   4.750   1.0714   0.01359   0.00618  -0.1099   0.0475   1.0000
   5.000   1.0934   0.01446   0.00692  -0.1089   0.0263   1.0000
   5.250   1.1156   0.01529   0.00777  -0.1079   0.0214   1.0000
   5.500   1.1392   0.01587   0.00842  -0.1071   0.0200   1.0000
   5.750   1.1618   0.01655   0.00919  -0.1062   0.0185   1.0000
   6.000   1.1834   0.01734   0.01003  -0.1052   0.0172   1.0000
   6.250   1.1995   0.01882   0.01160  -0.1032   0.0156   1.0000
   6.500   1.2199   0.01972   0.01259  -0.1020   0.0149   1.0000
   6.750   1.2399   0.02066   0.01363  -0.1006   0.0144   1.0000
   7.000   1.2592   0.02174   0.01480  -0.0992   0.0138   1.0000
   7.250   1.2779   0.02291   0.01608  -0.0978   0.0133   1.0000
   7.500   1.2965   0.02417   0.01743  -0.0963   0.0128   1.0000
   7.750   1.3147   0.02554   0.01890  -0.0948   0.0124   1.0000
   8.000   1.3324   0.02706   0.02054  -0.0933   0.0120   1.0000
   8.250   1.3492   0.02880   0.02237  -0.0918   0.0116   1.0000
   8.500   1.3627   0.03249   0.02627  -0.0900   0.0111   1.0000
   8.750   1.3746   0.03492   0.02901  -0.0878   0.0108   1.0000
   9.000   1.3861   0.03694   0.03130  -0.0855   0.0106   1.0000
   9.250   1.3939   0.03946   0.03411  -0.0828   0.0106   1.0000
   9.500   1.3976   0.04215   0.03710  -0.0797   0.0105   1.0000
   9.750   1.3952   0.04528   0.04052  -0.0760   0.0105   1.0000
  10.000   1.3853   0.04838   0.04389  -0.0713   0.0105   1.0000
  10.250   1.3705   0.05178   0.04755  -0.0666   0.0106   1.0000
  10.500   1.2597   0.04795   0.04393  -0.0501   0.0110   1.0000
  10.750   1.2392   0.05191   0.04810  -0.0469   0.0110   1.0000
  11.000   1.2222   0.05559   0.05194  -0.0446   0.0111   1.0000
  11.250   1.2051   0.05948   0.05599  -0.0430   0.0111   1.0000
  11.500   1.1841   0.06428   0.06096  -0.0420   0.0112   1.0000
  11.750   1.1598   0.06986   0.06671  -0.0417   0.0112   1.0000
  12.000   1.1391   0.07508   0.07208  -0.0421   0.0113   1.0000
<< Back to A18 (original) (a18-il)

Polar data table (+)

Polar graphs


<< Back to A18 (original) (a18-il)