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HAM-STD HS1-606 AIRFOIL (hs1606-il) Xfoil prediction polar at RE=200,000 Ncrit=9


Details Polar file
Airfoil: HAM-STD HS1-606 AIRFOIL (hs1606-il)
Reynolds number: 200,000
Max Cl/Cd: 95.44 at α=4.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-hs1606-il-200000.txt
Download as CSV file: xf-hs1606-il-200000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: HAM-STD HS1-606 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
  ------ -------- --------- --------- -------- -------- --------
  -7.500  -0.3795   0.09487   0.09166  -0.0234   1.0000   0.0337
  -7.250  -0.3951   0.09416   0.09103  -0.0194   1.0000   0.0339
  -7.000  -0.4051   0.09287   0.08980  -0.0169   1.0000   0.0344
  -6.750  -0.4120   0.09124   0.08822  -0.0152   1.0000   0.0350
  -6.500  -0.3955   0.08752   0.08449  -0.0196   0.9973   0.0364
  -6.250  -0.3694   0.08293   0.07988  -0.0272   0.9938   0.0387
  -6.000  -0.3203   0.07731   0.07417  -0.0464   0.9877   0.0425
  -5.750  -0.2764   0.07162   0.06836  -0.0589   0.9844   0.0429
  -5.500  -0.2678   0.06580   0.06260  -0.0589   0.9821   0.0444
  -5.250  -0.2474   0.06252   0.05929  -0.0608   0.9778   0.0459
  -5.000  -0.2156   0.05844   0.05514  -0.0665   0.9746   0.0483
  -4.750  -0.1441   0.05131   0.04761  -0.0843   0.9721   0.0551
  -4.500  -0.1009   0.03791   0.03389  -0.0941   0.9711   0.0299
  -4.250  -0.0467   0.02731   0.02241  -0.1043   0.9708   0.0283
  -4.000  -0.0005   0.02126   0.01519  -0.1093   0.9708   0.0262
  -3.750   0.0399   0.01866   0.01193  -0.1118   0.9703   0.0256
  -3.500   0.0735   0.01730   0.01029  -0.1129   0.9681   0.0262
  -3.250   0.1059   0.01641   0.00922  -0.1138   0.9654   0.0280
  -3.000   0.1415   0.01563   0.00827  -0.1153   0.9631   0.0336
  -2.750   0.1778   0.01506   0.00787  -0.1170   0.9610   0.0984
  -2.500   0.2123   0.01566   0.00842  -0.1187   0.9581   0.1299
  -2.250   0.2470   0.01583   0.00848  -0.1202   0.9551   0.1490
  -2.000   0.2769   0.01565   0.00826  -0.1207   0.9496   0.1629
  -1.750   0.3149   0.01535   0.00797  -0.1228   0.9465   0.1803
  -1.500   0.3542   0.01515   0.00776  -0.1253   0.9443   0.2003
  -1.250   0.3945   0.01493   0.00748  -0.1279   0.9428   0.2192
  -1.000   0.4181   0.01490   0.00750  -0.1272   0.9361   0.2334
  -0.750   0.4579   0.01450   0.00717  -0.1296   0.9327   0.2515
  -0.500   0.5008   0.01397   0.00670  -0.1324   0.9297   0.2715
  -0.250   0.5288   0.01372   0.00654  -0.1323   0.9216   0.2893
   0.000   0.5672   0.01315   0.00611  -0.1340   0.9173   0.3144
   0.250   0.5944   0.01287   0.00604  -0.1338   0.9095   0.3495
   0.500   0.6200   0.01129   0.00589  -0.1327   0.9052   1.0000
   0.750   0.6473   0.01138   0.00588  -0.1323   0.8986   1.0000
   1.000   0.6768   0.01134   0.00579  -0.1323   0.8927   1.0000
   1.250   0.7046   0.01136   0.00580  -0.1319   0.8861   1.0000
   1.500   0.7327   0.01132   0.00574  -0.1315   0.8789   1.0000
   1.750   0.7595   0.01133   0.00575  -0.1309   0.8709   1.0000
   2.000   0.7877   0.01125   0.00567  -0.1305   0.8635   1.0000
   2.250   0.8133   0.01127   0.00571  -0.1296   0.8535   1.0000
   2.500   0.8398   0.01121   0.00572  -0.1288   0.8437   1.0000
   2.750   0.8671   0.01111   0.00563  -0.1280   0.8338   1.0000
   3.000   0.8928   0.01106   0.00563  -0.1270   0.8216   1.0000
   3.250   0.9179   0.01093   0.00552  -0.1256   0.8050   1.0000
   3.500   0.9418   0.01069   0.00531  -0.1238   0.7800   1.0000
   3.750   0.9655   0.01054   0.00516  -0.1221   0.7494   1.0000
   4.000   0.9892   0.01048   0.00508  -0.1206   0.7100   1.0000
   4.250   1.0107   0.01059   0.00493  -0.1185   0.6295   1.0000
   4.500   1.0275   0.01142   0.00510  -0.1159   0.5215   1.0000
   4.750   1.0452   0.01236   0.00565  -0.1140   0.4375   1.0000
   5.000   1.0622   0.01340   0.00627  -0.1121   0.3568   1.0000
   5.250   1.0745   0.01501   0.00712  -0.1097   0.2163   1.0000
   5.500   1.0829   0.01747   0.00856  -0.1069   0.0813   1.0000
   5.750   1.1018   0.01859   0.00967  -0.1053   0.0679   1.0000
   6.000   1.1194   0.01984   0.01095  -0.1035   0.0615   1.0000
   6.250   1.1387   0.02089   0.01213  -0.1019   0.0571   1.0000
   6.500   1.1564   0.02223   0.01347  -0.1002   0.0533   1.0000
   6.750   1.1748   0.02394   0.01521  -0.0986   0.0498   1.0000
   7.000   1.1969   0.02513   0.01652  -0.0974   0.0472   1.0000
   7.250   1.2199   0.02663   0.01813  -0.0964   0.0451   1.0000
   7.500   1.2436   0.02831   0.01993  -0.0956   0.0432   1.0000
   7.750   1.2677   0.03037   0.02208  -0.0950   0.0412   1.0000
   8.000   1.2914   0.03424   0.02625  -0.0945   0.0384   1.0000
   8.250   1.3116   0.03610   0.02848  -0.0928   0.0371   1.0000
   8.500   1.3288   0.03762   0.03034  -0.0909   0.0338   1.0000
   8.750   1.3443   0.03894   0.03173  -0.0895   0.0301   1.0000
   9.000   1.3510   0.04300   0.03620  -0.0869   0.0267   1.0000
   9.250   1.3586   0.04571   0.03943  -0.0835   0.0243   1.0000
   9.500   1.3636   0.04851   0.04258  -0.0804   0.0220   1.0000
   9.750   1.2823   0.04432   0.03908  -0.0650   0.0226   1.0000
  10.250   1.3480   0.05727   0.05210  -0.0698   0.0175   1.0000
  10.500   1.3324   0.06051   0.05566  -0.0650   0.0171   1.0000
  10.750   1.3137   0.06429   0.05974  -0.0610   0.0169   1.0000
  11.000   1.2935   0.06846   0.06418  -0.0580   0.0169   1.0000
  11.250   1.2728   0.07304   0.06899  -0.0562   0.0170   1.0000
  11.500   1.2511   0.07810   0.07427  -0.0555   0.0170   1.0000
  11.750   1.2290   0.08370   0.08007  -0.0560   0.0171   1.0000
  12.000   1.2070   0.08990   0.08644  -0.0578   0.0173   1.0000
  12.250   1.1852   0.09684   0.09353  -0.0609   0.0176   1.0000
  12.500   1.1637   0.10474   0.10156  -0.0654   0.0179   1.0000
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