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

USA 31 AIRFOIL (usa31-il) Xfoil prediction polar at RE=200,000 Ncrit=9


Details Polar file
Airfoil: USA 31 AIRFOIL (usa31-il)
Reynolds number: 200,000
Max Cl/Cd: 62.3 at α=7.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-usa31-il-200000.txt
Download as CSV file: xf-usa31-il-200000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: USA 31 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
  ------ -------- --------- --------- -------- -------- --------
  -9.500   0.1558   0.10043   0.09686  -0.1111   0.8882   0.0541
  -9.250   0.1648   0.09695   0.09336  -0.1148   0.8863   0.0564
  -9.000   0.1382   0.09629   0.09273  -0.1153   0.8769   0.0571
  -8.750   0.1628   0.09159   0.08803  -0.1157   0.8744   0.0578
  -8.500   0.1888   0.08782   0.08425  -0.1172   0.8728   0.0590
  -8.250   0.2079   0.08433   0.08074  -0.1197   0.8711   0.0607
  -8.000   0.2237   0.08062   0.07701  -0.1233   0.8697   0.0635
  -7.750   0.1893   0.08032   0.07676  -0.1221   0.8569   0.0652
  -7.500   0.2168   0.07540   0.07183  -0.1236   0.8560   0.0665
  -7.250   0.2459   0.07169   0.06811  -0.1258   0.8551   0.0682
  -7.000   0.2673   0.06792   0.06431  -0.1293   0.8541   0.0709
  -6.750   0.1685   0.08203   0.07833  -0.1231   0.8549   0.0669
  -6.500   0.1963   0.07888   0.07516  -0.1257   0.8537   0.0688
  -5.750   0.1850   0.07148   0.06774  -0.1282   0.8258   0.0755
  -5.500   0.2137   0.06876   0.06501  -0.1279   0.8247   0.0770
  -5.250   0.2444   0.06578   0.06200  -0.1310   0.8232   0.0800
  -5.000   0.2605   0.06124   0.05735  -0.1394   0.8168   0.0865
  -4.750   0.2676   0.05978   0.05591  -0.1360   0.8090   0.0875
  -4.500   0.3023   0.05726   0.05335  -0.1386   0.8073   0.0910
  -4.250   0.3439   0.05262   0.04858  -0.1481   0.8051   0.0996
  -4.000   0.3684   0.05103   0.04696  -0.1482   0.8003   0.1031
  -3.000   0.4592   0.04225   0.03788  -0.1541   0.7740   0.1299
  -2.750   0.5075   0.03914   0.03458  -0.1609   0.7707   0.1433
  -2.500   0.4831   0.02343   0.01643  -0.1569   0.7597   0.0652
  -2.250   0.5214   0.02158   0.01437  -0.1588   0.7555   0.0636
  -2.000   0.5713   0.02015   0.01264  -0.1627   0.7524   0.0621
  -1.750   0.5717   0.01992   0.01232  -0.1570   0.7423   0.0623
  -1.500   0.6131   0.01910   0.01131  -0.1593   0.7374   0.0633
  -1.250   0.6455   0.01849   0.01059  -0.1598   0.7313   0.0635
  -1.000   0.6606   0.01815   0.01020  -0.1570   0.7228   0.0635
  -0.750   0.7057   0.01752   0.00948  -0.1601   0.7180   0.0639
  -0.500   0.7136   0.01742   0.00937  -0.1558   0.7089   0.0643
  -0.250   0.7479   0.01707   0.00896  -0.1568   0.7025   0.0653
   0.000   0.7701   0.01688   0.00877  -0.1554   0.6950   0.0667
   0.250   0.7926   0.01676   0.00861  -0.1541   0.6872   0.0686
   0.500   0.8337   0.01656   0.00828  -0.1566   0.6811   0.0744
   0.750   0.8427   0.01652   0.00826  -0.1526   0.6720   0.0805
   1.000   0.8852   0.01609   0.00828  -0.1556   0.6656   0.2286
   1.250   0.8928   0.01652   0.00870  -0.1512   0.6567   0.2687
   1.500   0.9223   0.01710   0.00917  -0.1512   0.6494   0.2931
   1.750   0.9381   0.01758   0.00957  -0.1486   0.6413   0.3074
   2.000   0.9589   0.01804   0.01003  -0.1469   0.6332   0.3208
   2.250   0.9845   0.01834   0.01029  -0.1465   0.6258   0.3325
   2.500   0.9989   0.01858   0.01043  -0.1437   0.6176   0.3404
   2.750   1.0386   0.01879   0.01051  -0.1461   0.6109   0.3557
   3.000   1.0387   0.01889   0.01072  -0.1407   0.6022   0.3618
   3.250   1.0673   0.01896   0.01070  -0.1409   0.5951   0.3768
   3.500   1.0812   0.01925   0.01093  -0.1382   0.5874   0.3849
   3.750   1.0999   0.01925   0.01092  -0.1366   0.5799   0.3896
   4.000   1.1296   0.01932   0.01089  -0.1371   0.5731   0.3940
   4.250   1.1369   0.01959   0.01117  -0.1332   0.5651   0.3961
   4.500   1.1651   0.01972   0.01116  -0.1334   0.5587   0.3994
   4.750   1.1788   0.01994   0.01140  -0.1309   0.5515   0.4009
   5.000   1.1957   0.02008   0.01153  -0.1289   0.5444   0.4024
   5.250   1.2285   0.02013   0.01147  -0.1301   0.5383   0.4046
   5.500   1.2331   0.02047   0.01190  -0.1259   0.5311   0.4058
   5.750   1.2545   0.02062   0.01201  -0.1249   0.5247   0.4086
   6.000   1.2780   0.02085   0.01218  -0.1243   0.5186   0.4115
   6.250   1.2875   0.02119   0.01256  -0.1212   0.5116   0.4130
   6.500   1.3128   0.02137   0.01267  -0.1210   0.5059   0.4147
   6.750   1.3298   0.02164   0.01298  -0.1193   0.4994   0.4165
   7.000   1.3433   0.02194   0.01332  -0.1171   0.4927   0.4185
   7.250   1.3737   0.02205   0.01334  -0.1178   0.4866   0.4212
   7.500   1.3796   0.02252   0.01391  -0.1143   0.4800   0.4230
   7.750   1.3954   0.02283   0.01422  -0.1125   0.4733   0.4254
   8.000   1.4251   0.02308   0.01439  -0.1133   0.4673   0.4288
   8.250   1.4366   0.02363   0.01508  -0.1112   0.4607   0.4319
   8.500   1.4616   0.02395   0.01541  -0.1114   0.4546   0.4355
   8.750   1.4880   0.02430   0.01574  -0.1117   0.4489   0.4391
   9.000   1.4967   0.02491   0.01645  -0.1090   0.4428   0.4418
   9.250   1.5165   0.02529   0.01683  -0.1082   0.4371   0.4453
   9.500   1.5408   0.02567   0.01721  -0.1082   0.4314   0.4496
   9.750   1.5467   0.02638   0.01803  -0.1052   0.4252   0.4529
  10.000   1.5653   0.02681   0.01845  -0.1043   0.4194   0.4578
  10.250   1.5816   0.02736   0.01902  -0.1030   0.4130   0.4635
  10.500   1.5875   0.02814   0.01993  -0.1003   0.4067   0.4708
  10.750   1.6080   0.02853   0.02032  -0.0998   0.4005   0.4880
  11.000   1.6752   0.02985   0.02249  -0.1112   0.3901   1.0000
  11.250   1.6880   0.03049   0.02310  -0.1094   0.3841   1.0000
  11.500   1.6989   0.03130   0.02392  -0.1075   0.3782   1.0000
  11.750   1.7001   0.03244   0.02517  -0.1044   0.3718   1.0000
  12.000   1.7139   0.03309   0.02578  -0.1029   0.3660   1.0000
  12.250   1.7172   0.03427   0.02706  -0.1003   0.3597   1.0000
  12.500   1.7188   0.03554   0.02841  -0.0975   0.3533   1.0000
  12.750   1.7365   0.03605   0.02880  -0.0966   0.3469   1.0000
  13.000   1.7266   0.03805   0.03100  -0.0929   0.3400   1.0000
  13.250   1.7311   0.03930   0.03226  -0.0908   0.3331   1.0000
  13.500   1.7334   0.04081   0.03382  -0.0886   0.3263   1.0000
  13.750   1.7306   0.04269   0.03578  -0.0861   0.3188   1.0000
  14.000   1.7350   0.04413   0.03719  -0.0842   0.3116   1.0000
  14.250   1.7293   0.04646   0.03965  -0.0819   0.3040   1.0000
  14.500   1.7359   0.04782   0.04092  -0.0804   0.2969   1.0000
  14.750   1.7285   0.05055   0.04383  -0.0783   0.2894   1.0000
  15.000   1.7313   0.05233   0.04557  -0.0768   0.2824   1.0000
  15.250   1.7257   0.05507   0.04844  -0.0751   0.2749   1.0000
  15.500   1.7250   0.05735   0.05072  -0.0737   0.2677   1.0000
  15.750   1.7237   0.05981   0.05324  -0.0725   0.2611   1.0000
  16.000   1.7195   0.06262   0.05612  -0.0712   0.2539   1.0000
  16.250   1.7195   0.06496   0.05846  -0.0702   0.2474   1.0000
  16.500   1.7122   0.06828   0.06190  -0.0692   0.2401   1.0000
  16.750   1.7123   0.07066   0.06420  -0.0683   0.2335   1.0000
  17.000   1.7032   0.07438   0.06809  -0.0675   0.2264   1.0000
  17.250   1.7008   0.07715   0.07083  -0.0668   0.2201   1.0000
  17.500   1.6937   0.08068   0.07449  -0.0663   0.2136   1.0000
  17.750   1.6890   0.08389   0.07774  -0.0658   0.2074   1.0000
  18.000   1.6850   0.08701   0.08087  -0.0654   0.2015   1.0000
  18.250   1.6771   0.09080   0.08479  -0.0652   0.1953   1.0000
  18.500   1.6745   0.09372   0.08764  -0.0649   0.1894   1.0000
  18.750   1.6651   0.09785   0.09194  -0.0650   0.1836   1.0000
  19.000   1.6592   0.10138   0.09550  -0.0651   0.1779   1.0000
  19.250   1.6545   0.10475   0.09888  -0.0652   0.1724   1.0000
<< Back to USA 31 AIRFOIL (usa31-il)

Polar data table (+)

Polar graphs


<< Back to USA 31 AIRFOIL (usa31-il)