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

USA 98 AIRFOIL (usa98-il) Xfoil prediction polar at RE=200,000 Ncrit=9


Details Polar file
Airfoil: USA 98 AIRFOIL (usa98-il)
Reynolds number: 200,000
Max Cl/Cd: 84.09 at α=7°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-usa98-il-200000.txt
Download as CSV file: xf-usa98-il-200000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: USA 98 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
  ------ -------- --------- --------- -------- -------- --------
  -8.250  -0.0968   0.10151   0.09759  -0.0769   0.9690   0.0766
  -8.000  -0.0818   0.09814   0.09421  -0.0807   0.9598   0.0787
  -7.750  -0.0894   0.09299   0.08905  -0.0922   0.9462   0.0813
  -7.500  -0.0582   0.08961   0.08567  -0.0930   0.9441   0.0819
  -7.250  -0.0286   0.08647   0.08252  -0.0954   0.9418   0.0829
  -7.000  -0.0113   0.08389   0.07994  -0.0968   0.9313   0.0844
  -6.750  -0.0152   0.07756   0.07356  -0.1157   0.9135   0.0903
  -6.500   0.0070   0.07488   0.07088  -0.1128   0.9084   0.0908
  -6.250   0.0306   0.07256   0.06855  -0.1123   0.9012   0.0916
  -6.000   0.0595   0.06992   0.06587  -0.1145   0.8963   0.0930
  -5.750   0.0801   0.06738   0.06330  -0.1172   0.8860   0.0951
  -5.500   0.1058   0.06019   0.05598  -0.1350   0.8754   0.1016
  -5.250   0.1497   0.05213   0.04749  -0.1586   0.8649   0.1123
  -5.000   0.1679   0.05059   0.04608  -0.1547   0.8573   0.1130
  -4.750   0.1863   0.04999   0.04551  -0.1514   0.8490   0.1138
  -4.500   0.2081   0.04913   0.04461  -0.1498   0.8418   0.1151
  -4.250   0.2406   0.04583   0.04117  -0.1575   0.8325   0.1236
  -4.000   0.2872   0.03965   0.03441  -0.1731   0.8254   0.1402
  -3.750   0.3043   0.03806   0.03299  -0.1711   0.8171   0.1415
  -3.500   0.3267   0.03731   0.03228  -0.1698   0.8098   0.1432
  -3.250   0.3931   0.02497   0.01816  -0.1874   0.8046   0.1022
  -3.000   0.4216   0.02342   0.01645  -0.1879   0.7966   0.0969
  -2.750   0.4572   0.02196   0.01425  -0.1893   0.7898   0.0904
  -2.500   0.4875   0.02114   0.01324  -0.1899   0.7836   0.0899
  -2.250   0.5156   0.02033   0.01232  -0.1900   0.7759   0.0898
  -2.000   0.5459   0.01976   0.01156  -0.1904   0.7693   0.0901
  -1.750   0.5753   0.01891   0.01064  -0.1908   0.7631   0.0915
  -1.500   0.6027   0.01843   0.01016  -0.1907   0.7557   0.0929
  -1.250   0.6323   0.01802   0.00968  -0.1909   0.7493   0.0940
  -1.000   0.6618   0.01771   0.00931  -0.1910   0.7433   0.0952
  -0.750   0.6889   0.01742   0.00903  -0.1907   0.7359   0.0966
  -0.500   0.7183   0.01715   0.00872  -0.1908   0.7297   0.0983
  -0.250   0.7482   0.01704   0.00853  -0.1909   0.7240   0.1005
   0.000   0.7748   0.01666   0.00830  -0.1907   0.7165   0.1042
   0.250   0.8040   0.01647   0.00813  -0.1908   0.7102   0.1081
   0.500   0.8347   0.01640   0.00798  -0.1911   0.7050   0.1121
   0.750   0.8613   0.01622   0.00793  -0.1909   0.6975   0.1182
   1.000   0.8906   0.01607   0.00779  -0.1910   0.6910   0.1298
   1.250   0.9227   0.01555   0.00779  -0.1920   0.6859   0.2946
   1.500   0.9481   0.01566   0.00817  -0.1914   0.6786   0.4048
   1.750   0.9758   0.01577   0.00831  -0.1911   0.6720   0.4420
   2.000   1.0054   0.01587   0.00838  -0.1911   0.6668   0.4701
   2.250   1.0305   0.01599   0.00860  -0.1905   0.6596   0.4926
   2.500   1.0577   0.01603   0.00868  -0.1902   0.6530   0.5170
   2.750   1.0870   0.01605   0.00870  -0.1902   0.6476   0.5464
   3.000   1.1114   0.01610   0.00893  -0.1894   0.6402   0.5806
   3.250   1.1380   0.01599   0.00894  -0.1889   0.6326   0.6208
   3.500   1.1641   0.01584   0.00894  -0.1882   0.6251   0.6823
   3.750   1.1832   0.01533   0.00884  -0.1859   0.6165   1.0000
   4.000   1.2129   0.01546   0.00880  -0.1861   0.6101   1.0000
   4.250   1.2381   0.01568   0.00901  -0.1855   0.6026   1.0000
   4.500   1.2649   0.01585   0.00912  -0.1851   0.5955   1.0000
   4.750   1.2937   0.01604   0.00919  -0.1851   0.5894   1.0000
   5.000   1.3171   0.01627   0.00946  -0.1842   0.5814   1.0000
   5.250   1.3442   0.01644   0.00958  -0.1838   0.5747   1.0000
   5.500   1.3700   0.01669   0.00980  -0.1833   0.5676   1.0000
   5.750   1.3941   0.01690   0.01002  -0.1825   0.5596   1.0000
   6.000   1.4225   0.01712   0.01013  -0.1824   0.5530   1.0000
   6.250   1.4438   0.01739   0.01049  -0.1811   0.5446   1.0000
   6.500   1.4692   0.01758   0.01061  -0.1804   0.5364   1.0000
   6.750   1.4906   0.01782   0.01089  -0.1790   0.5268   1.0000
   7.000   1.5145   0.01801   0.01101  -0.1781   0.5172   1.0000
   7.250   1.5341   0.01828   0.01134  -0.1764   0.5070   1.0000
   7.500   1.5577   0.01853   0.01150  -0.1754   0.4978   1.0000
   7.750   1.5748   0.01882   0.01189  -0.1732   0.4869   1.0000
   8.000   1.5956   0.01911   0.01212  -0.1718   0.4767   1.0000
   8.250   1.6118   0.01942   0.01248  -0.1694   0.4652   1.0000
   8.500   1.6282   0.01977   0.01285  -0.1672   0.4538   1.0000
   8.750   1.6445   0.02015   0.01316  -0.1649   0.4425   1.0000
   9.000   1.6555   0.02056   0.01365  -0.1617   0.4297   1.0000
   9.250   1.6647   0.02102   0.01412  -0.1582   0.4168   1.0000
   9.500   1.6731   0.02159   0.01464  -0.1546   0.4041   1.0000
   9.750   1.6804   0.02223   0.01529  -0.1509   0.3904   1.0000
  10.000   1.6867   0.02297   0.01604  -0.1473   0.3756   1.0000
  10.250   1.6929   0.02382   0.01687  -0.1437   0.3610   1.0000
  10.500   1.6986   0.02477   0.01781  -0.1403   0.3471   1.0000
  10.750   1.7033   0.02585   0.01885  -0.1369   0.3336   1.0000
  11.000   1.7069   0.02706   0.02004  -0.1335   0.3200   1.0000
  11.250   1.7112   0.02833   0.02133  -0.1304   0.3063   1.0000
  11.500   1.7146   0.02975   0.02274  -0.1274   0.2935   1.0000
  11.750   1.7163   0.03136   0.02434  -0.1245   0.2813   1.0000
  12.000   1.7167   0.03318   0.02612  -0.1216   0.2689   1.0000
  12.250   1.7185   0.03502   0.02801  -0.1192   0.2561   1.0000
  12.500   1.7184   0.03710   0.03010  -0.1168   0.2435   1.0000
  12.750   1.7163   0.03945   0.03244  -0.1145   0.2306   1.0000
  13.000   1.7123   0.04209   0.03506  -0.1122   0.2165   1.0000
  13.250   1.7066   0.04501   0.03797  -0.1102   0.2004   1.0000
  13.500   1.6986   0.04832   0.04126  -0.1082   0.1809   1.0000
  13.750   1.6852   0.05231   0.04516  -0.1063   0.1599   1.0000
  14.000   1.6709   0.05653   0.04929  -0.1046   0.1384   1.0000
  14.250   1.6560   0.06097   0.05364  -0.1031   0.1231   1.0000
  14.500   1.6439   0.06524   0.05787  -0.1019   0.1128   1.0000
  14.750   1.6337   0.06937   0.06200  -0.1009   0.1056   1.0000
  15.000   1.6235   0.07360   0.06621  -0.1001   0.1003   1.0000
  15.250   1.6178   0.07734   0.07000  -0.0994   0.0958   1.0000
  15.500   1.6118   0.08116   0.07384  -0.0989   0.0921   1.0000
  15.750   1.6044   0.08508   0.07773  -0.0983   0.0890   1.0000
  16.000   1.6036   0.08829   0.08104  -0.0979   0.0859   1.0000
  16.250   1.6016   0.09161   0.08438  -0.0976   0.0831   1.0000
  16.500   1.5994   0.09472   0.08742  -0.0969   0.0806   1.0000
<< Back to USA 98 AIRFOIL (usa98-il)

Polar data table (+)

Polar graphs


<< Back to USA 98 AIRFOIL (usa98-il)