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A18 (original) (a18-il) Xfoil prediction polar at RE=50,000 Ncrit=5


Details Polar file
Airfoil: A18 (original) (a18-il)
Reynolds number: 50,000
Max Cl/Cd: 44.62 at α=4.75°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-a18-il-50000-n5.txt
Download as CSV file: xf-a18-il-50000-n5.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 =     0.050 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.000  -0.4071   0.09802   0.09099  -0.0289   1.0000   0.0592
  -8.750  -0.4102   0.09505   0.08811  -0.0290   1.0000   0.0593
  -8.500  -0.4116   0.09188   0.08503  -0.0299   1.0000   0.0602
  -8.250  -0.4115   0.08845   0.08169  -0.0316   1.0000   0.0612
  -8.000  -0.4099   0.08470   0.07801  -0.0339   1.0000   0.0620
  -7.750  -0.4063   0.08056   0.07392  -0.0369   1.0000   0.0626
  -7.500  -0.4002   0.07601   0.06941  -0.0404   1.0000   0.0627
  -7.250  -0.3908   0.07090   0.06429  -0.0449   1.0000   0.0629
  -7.000  -0.3763   0.06468   0.05799  -0.0514   1.0000   0.0638
  -6.750  -0.3539   0.05674   0.04982  -0.0607   1.0000   0.0674
  -6.500  -0.3406   0.05568   0.04875  -0.0592   1.0000   0.0729
  -6.250  -0.3105   0.04826   0.04089  -0.0674   1.0000   0.0761
  -6.000  -0.2827   0.04384   0.03607  -0.0720   1.0000   0.0844
  -5.750  -0.2518   0.03932   0.03099  -0.0762   1.0000   0.0892
  -5.500  -0.2160   0.03458   0.02532  -0.0809   1.0000   0.0950
  -5.250  -0.1943   0.03373   0.02442  -0.0805   1.0000   0.1019
  -5.000  -0.1630   0.03102   0.02104  -0.0825   1.0000   0.1057
  -4.750  -0.1328   0.02887   0.01828  -0.0838   1.0000   0.1088
  -4.500  -0.1070   0.02768   0.01693  -0.0840   1.0000   0.1122
  -4.250  -0.0806   0.02674   0.01573  -0.0842   1.0000   0.1190
  -4.000  -0.0548   0.02590   0.01464  -0.0842   1.0000   0.1261
  -3.750  -0.0286   0.02513   0.01369  -0.0843   1.0000   0.1319
  -3.500  -0.0016   0.02439   0.01267  -0.0843   1.0000   0.1383
  -3.250   0.0245   0.02388   0.01209  -0.0844   1.0000   0.1465
  -3.000   0.0511   0.02355   0.01167  -0.0845   1.0000   0.1609
  -2.750   0.0778   0.02324   0.01129  -0.0847   1.0000   0.1787
  -2.500   0.1048   0.02293   0.01101  -0.0850   1.0000   0.2006
  -2.250   0.1317   0.02272   0.01081  -0.0853   1.0000   0.2380
  -2.000   0.1632   0.02244   0.01082  -0.0868   0.9975   0.2925
  -1.750   0.2030   0.02223   0.01088  -0.0897   0.9904   0.3841
  -1.500   0.2417   0.02162   0.01089  -0.0923   0.9840   0.5131
  -1.250   0.2670   0.02038   0.01064  -0.0913   0.9711   1.0000
  -1.000   0.3075   0.02063   0.01054  -0.0942   0.9611   1.0000
  -0.750   0.3487   0.02087   0.01052  -0.0971   0.9516   1.0000
  -0.500   0.3849   0.02107   0.01056  -0.0991   0.9401   1.0000
  -0.250   0.4207   0.02127   0.01062  -0.1009   0.9284   1.0000
   0.000   0.4573   0.02143   0.01069  -0.1028   0.9169   1.0000
   0.250   0.4954   0.02156   0.01077  -0.1049   0.9063   1.0000
   0.500   0.5336   0.02166   0.01085  -0.1069   0.8960   1.0000
   1.000   0.5996   0.02199   0.01123  -0.1090   0.8723   1.0000
   1.250   0.6326   0.02214   0.01143  -0.1100   0.8610   1.0000
   1.500   0.6664   0.02226   0.01161  -0.1110   0.8500   1.0000
   1.750   0.7008   0.02233   0.01181  -0.1120   0.8391   1.0000
   2.000   0.7313   0.02251   0.01210  -0.1123   0.8262   1.0000
   2.250   0.7625   0.02261   0.01232  -0.1126   0.8123   1.0000
   2.500   0.7955   0.02249   0.01236  -0.1128   0.7962   1.0000
   2.750   0.8231   0.02246   0.01247  -0.1119   0.7748   1.0000
   3.000   0.8547   0.02222   0.01236  -0.1114   0.7537   1.0000
   3.250   0.8817   0.02224   0.01255  -0.1103   0.7305   1.0000
   3.500   0.9104   0.02218   0.01263  -0.1094   0.7067   1.0000
   3.750   0.9367   0.02223   0.01283  -0.1082   0.6802   1.0000
   4.000   0.9616   0.02243   0.01320  -0.1069   0.6535   1.0000
   4.250   0.9866   0.02262   0.01360  -0.1056   0.6240   1.0000
   4.500   1.0105   0.02284   0.01395  -0.1039   0.5892   1.0000
   4.750   1.0324   0.02314   0.01427  -0.1018   0.5429   1.0000
   5.000   1.0494   0.02369   0.01478  -0.0990   0.4771   1.0000
   5.250   1.0632   0.02458   0.01526  -0.0959   0.3944   1.0000
   5.500   1.0745   0.02598   0.01632  -0.0931   0.2924   1.0000
   5.750   1.0810   0.02823   0.01777  -0.0904   0.2119   1.0000
   6.000   1.0911   0.03036   0.01950  -0.0884   0.1428   1.0000
   6.250   1.1017   0.03253   0.02129  -0.0863   0.0921   1.0000
   6.500   1.1125   0.03480   0.02344  -0.0839   0.0702   1.0000
   6.750   1.1250   0.03691   0.02560  -0.0817   0.0585   1.0000
   7.000   1.1382   0.03909   0.02782  -0.0797   0.0513   1.0000
   7.250   1.1571   0.04122   0.03022  -0.0779   0.0467   1.0000
   7.500   1.1744   0.04346   0.03252  -0.0765   0.0425   1.0000
   7.750   1.1953   0.04601   0.03544  -0.0754   0.0388   1.0000
   8.000   1.2160   0.04882   0.03861  -0.0742   0.0364   1.0000
   8.250   1.2334   0.05193   0.04206  -0.0729   0.0350   1.0000
   8.500   1.2462   0.05511   0.04553  -0.0713   0.0337   1.0000
   8.750   1.2549   0.05844   0.04906  -0.0697   0.0324   1.0000
   9.000   1.2566   0.06214   0.05319  -0.0672   0.0316   1.0000
   9.250   1.2519   0.06587   0.05741  -0.0642   0.0310   1.0000
   9.500   1.2410   0.06949   0.06142  -0.0610   0.0306   1.0000
   9.750   1.2269   0.07336   0.06564  -0.0582   0.0304   1.0000
  10.000   1.2106   0.07758   0.07017  -0.0563   0.0303   1.0000
  10.250   1.1925   0.08227   0.07514  -0.0554   0.0303   1.0000
  10.500   1.1731   0.08749   0.08060  -0.0556   0.0304   1.0000
  10.750   1.1528   0.09337   0.08670  -0.0571   0.0305   1.0000
  11.000   1.1327   0.09993   0.09343  -0.0598   0.0308   1.0000
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