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EPPLER 422 AIRFOIL (e422-il) Xfoil prediction polar at RE=100,000 Ncrit=5


Details Polar file
Airfoil: EPPLER 422 AIRFOIL (e422-il)
Reynolds number: 100,000
Max Cl/Cd: 46.71 at α=6°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-e422-il-100000-n5.txt
Download as CSV file: xf-e422-il-100000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: EPPLER 422 AIRFOIL                              
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.100 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.500   0.0391   0.08810   0.08223  -0.0873   0.7118   0.0305
  -8.250   0.0418   0.08465   0.07875  -0.0891   0.7010   0.0302
  -8.000   0.0532   0.08259   0.07664  -0.0895   0.6900   0.0296
  -7.750   0.0572   0.07958   0.07360  -0.0908   0.6805   0.0297
  -7.500   0.0611   0.07668   0.07069  -0.0920   0.6719   0.0297
  -7.250   0.0627   0.07367   0.06768  -0.0931   0.6632   0.0298
  -7.000   0.0624   0.07077   0.06476  -0.0940   0.6560   0.0299
  -6.750   0.0643   0.06809   0.06210  -0.0951   0.6481   0.0297
  -6.500   0.0675   0.06463   0.05860  -0.0974   0.6417   0.0296
  -6.250   0.0741   0.06135   0.05530  -0.0999   0.6351   0.0292
  -6.000   0.0806   0.05702   0.05094  -0.1036   0.6288   0.0289
  -5.750   0.0882   0.05104   0.04485  -0.1092   0.6239   0.0285
  -5.500   0.0896   0.03387   0.02660  -0.1242   0.6221   0.0271
  -5.250   0.1092   0.03063   0.02282  -0.1254   0.6156   0.0275
  -5.000   0.1311   0.02840   0.02013  -0.1256   0.6090   0.0282
  -4.750   0.1546   0.02665   0.01789  -0.1254   0.6035   0.0290
  -4.500   0.1786   0.02521   0.01605  -0.1250   0.5970   0.0301
  -4.250   0.2030   0.02439   0.01513  -0.1246   0.5907   0.0317
  -4.000   0.2287   0.02355   0.01396  -0.1242   0.5856   0.0348
  -3.750   0.2538   0.02293   0.01327  -0.1238   0.5801   0.0377
  -3.500   0.2791   0.02224   0.01243  -0.1234   0.5745   0.0412
  -3.250   0.3052   0.02165   0.01168  -0.1230   0.5695   0.0467
  -3.000   0.3316   0.02112   0.01103  -0.1226   0.5651   0.0560
  -2.750   0.3573   0.02062   0.01055  -0.1223   0.5595   0.0729
  -2.500   0.3838   0.02033   0.01031  -0.1220   0.5545   0.1005
  -2.250   0.4111   0.02037   0.01028  -0.1217   0.5504   0.1296
  -2.000   0.4383   0.02051   0.01028  -0.1215   0.5465   0.1534
  -1.750   0.4643   0.02063   0.01035  -0.1210   0.5416   0.1714
  -1.500   0.4909   0.02069   0.01029  -0.1206   0.5371   0.1873
  -1.250   0.5176   0.02068   0.01020  -0.1203   0.5332   0.2004
  -1.000   0.5449   0.02065   0.01005  -0.1200   0.5298   0.2126
  -0.750   0.5705   0.02068   0.01002  -0.1195   0.5254   0.2247
  -0.500   0.5966   0.02069   0.01004  -0.1191   0.5215   0.2361
  -0.250   0.6232   0.02070   0.00998  -0.1188   0.5179   0.2481
   0.000   0.6506   0.02072   0.00989  -0.1186   0.5147   0.2614
   0.250   0.6774   0.02075   0.00988  -0.1183   0.5115   0.2748
   0.500   0.7022   0.02085   0.01004  -0.1177   0.5074   0.2887
   0.750   0.7284   0.02095   0.01014  -0.1174   0.5036   0.3051
   1.000   0.7553   0.02104   0.01023  -0.1171   0.5004   0.3240
   1.250   0.7827   0.02112   0.01029  -0.1170   0.4976   0.3471
   1.500   0.8105   0.02118   0.01033  -0.1169   0.4951   0.3768
   1.750   0.8342   0.02132   0.01063  -0.1163   0.4914   0.4159
   2.000   0.8574   0.02133   0.01092  -0.1155   0.4877   0.4890
   2.500   0.9240   0.02102   0.01125  -0.1176   0.4814   1.0000
   2.750   0.9506   0.02132   0.01139  -0.1173   0.4790   1.0000
   3.000   0.9763   0.02166   0.01161  -0.1170   0.4765   1.0000
   3.250   0.9971   0.02215   0.01213  -0.1159   0.4728   1.0000
   3.500   1.0195   0.02257   0.01254  -0.1151   0.4694   1.0000
   3.750   1.0432   0.02296   0.01288  -0.1144   0.4664   1.0000
   4.000   1.0681   0.02332   0.01317  -0.1139   0.4639   1.0000
   4.250   1.0942   0.02364   0.01340  -0.1136   0.4617   1.0000
   4.500   1.1194   0.02403   0.01373  -0.1132   0.4594   1.0000
   4.750   1.1365   0.02468   0.01448  -0.1117   0.4556   1.0000
   5.000   1.1564   0.02523   0.01507  -0.1106   0.4522   1.0000
   5.250   1.1783   0.02570   0.01555  -0.1098   0.4494   1.0000
   5.500   1.2019   0.02612   0.01595  -0.1092   0.4470   1.0000
   5.750   1.2276   0.02646   0.01625  -0.1089   0.4447   1.0000
   6.000   1.2533   0.02683   0.01656  -0.1087   0.4425   1.0000
   6.250   1.2639   0.02776   0.01767  -0.1064   0.4386   1.0000
   6.500   1.2797   0.02851   0.01851  -0.1048   0.4353   1.0000
   6.750   1.2987   0.02910   0.01916  -0.1037   0.4324   1.0000
   7.000   1.3210   0.02952   0.01960  -0.1030   0.4298   1.0000
   7.250   1.3472   0.02982   0.01985  -0.1028   0.4276   1.0000
   7.500   1.3688   0.03035   0.02042  -0.1021   0.4252   1.0000
   7.750   1.3682   0.03174   0.02204  -0.0985   0.4209   1.0000
   8.000   1.3776   0.03270   0.02310  -0.0963   0.4173   1.0000
   8.250   1.3948   0.03327   0.02372  -0.0950   0.4143   1.0000
   8.500   1.4187   0.03357   0.02403  -0.0945   0.4119   1.0000
   8.750   1.4495   0.03365   0.02408  -0.0949   0.4098   1.0000
   9.000   1.4203   0.03601   0.02671  -0.0880   0.4047   1.0000
   9.250   1.4082   0.03783   0.02866  -0.0836   0.4002   1.0000
   9.500   1.4234   0.03843   0.02930  -0.0823   0.3972   1.0000
   9.750   1.4537   0.03828   0.02917  -0.0824   0.3948   1.0000
  10.000   1.2178   0.06080   0.05213  -0.0727   0.3707   1.0000
  10.250   1.2749   0.05668   0.04803  -0.0717   0.3733   1.0000
  10.500   1.3318   0.05301   0.04436  -0.0711   0.3746   1.0000
  10.750   1.4019   0.04859   0.03992  -0.0714   0.3760   1.0000
  13.750   1.1994   0.10890   0.10120  -0.0733   0.2760   1.0000
  14.000   1.2219   0.10843   0.10079  -0.0726   0.2742   1.0000
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