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USA 98 AIRFOIL (usa98-il) Xfoil prediction polar at RE=50,000 Ncrit=5


Details Polar file
Airfoil: USA 98 AIRFOIL (usa98-il)
Reynolds number: 50,000
Max Cl/Cd: 26.56 at α=6°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-usa98-il-50000-n5.txt
Download as CSV file: xf-usa98-il-50000-n5.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 =     0.050 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -7.250  -0.1477   0.11311   0.10571  -0.0579   0.9473   0.1670
  -7.000  -0.1373   0.11077   0.10337  -0.0598   0.9382   0.1724
  -6.750  -0.1607   0.11134   0.10400  -0.0634   0.9230   0.1772
  -6.500  -0.1264   0.10600   0.09864  -0.0644   0.9197   0.1800
  -6.250  -0.1043   0.10280   0.09543  -0.0655   0.9134   0.1843
  -6.000  -0.0968   0.10067   0.09331  -0.0668   0.9043   0.1893
  -5.250  -0.0849   0.08355   0.07586  -0.0913   0.8728   0.1148
  -5.000  -0.0585   0.08084   0.07314  -0.0917   0.8693   0.1132
  -4.750  -0.0489   0.07837   0.07067  -0.0925   0.8593   0.1120
  -4.500  -0.0225   0.07429   0.06652  -0.0983   0.8540   0.1109
  -4.250   0.0013   0.07033   0.06247  -0.1038   0.8475   0.1097
  -4.000   0.0467   0.05800   0.04945  -0.1267   0.8394   0.1021
  -3.750   0.0873   0.05395   0.04517  -0.1333   0.8355   0.1015
  -3.500   0.1375   0.04927   0.04006  -0.1424   0.8328   0.1012
  -3.250   0.1658   0.04635   0.03662  -0.1468   0.8234   0.1020
  -3.000   0.2035   0.04431   0.03446  -0.1502   0.8186   0.1036
  -2.750   0.2478   0.04210   0.03198  -0.1547   0.8153   0.1049
  -2.500   0.2793   0.04059   0.03018  -0.1568   0.8082   0.1054
  -2.250   0.3116   0.03931   0.02862  -0.1587   0.8014   0.1062
  -2.000   0.3529   0.03783   0.02685  -0.1617   0.7972   0.1074
  -1.750   0.3970   0.03648   0.02522  -0.1649   0.7940   0.1102
  -1.500   0.4136   0.03647   0.02504  -0.1636   0.7834   0.1124
  -1.250   0.4502   0.03567   0.02396  -0.1652   0.7783   0.1151
  -1.000   0.4900   0.03467   0.02292  -0.1673   0.7746   0.1176
  -0.750   0.5063   0.03489   0.02313  -0.1657   0.7648   0.1194
  -0.500   0.5385   0.03447   0.02263  -0.1664   0.7589   0.1224
  -0.250   0.5778   0.03386   0.02186  -0.1681   0.7549   0.1278
   0.000   0.5951   0.03423   0.02226  -0.1668   0.7457   0.1321
   0.250   0.6251   0.03409   0.02210  -0.1672   0.7391   0.1386
   0.500   0.6642   0.03359   0.02153  -0.1688   0.7350   0.1465
   0.750   0.6821   0.03409   0.02209  -0.1676   0.7261   0.1556
   1.000   0.7108   0.03407   0.02220  -0.1680   0.7193   0.1761
   1.250   0.7499   0.03346   0.02216  -0.1698   0.7152   0.2737
   1.500   0.7647   0.03433   0.02324  -0.1679   0.7063   0.3677
   1.750   0.7879   0.03480   0.02386  -0.1670   0.6992   0.4430
   2.000   0.8221   0.03466   0.02381  -0.1674   0.6950   0.5117
   2.250   0.8289   0.03575   0.02512  -0.1643   0.6852   0.5560
   2.500   0.8517   0.03586   0.02546  -0.1630   0.6790   0.6172
   2.750   0.8860   0.03542   0.02513  -0.1634   0.6751   0.6660
   3.000   0.8877   0.03676   0.02665  -0.1598   0.6643   0.6994
   3.250   0.9053   0.03638   0.02659  -0.1574   0.6585   0.9053
   3.500   0.9479   0.03617   0.02612  -0.1595   0.6549   1.0000
   3.750   0.9429   0.03833   0.02827  -0.1556   0.6425   1.0000
   4.000   0.9786   0.03834   0.02810  -0.1565   0.6380   1.0000
   4.500   1.0059   0.04054   0.03017  -0.1528   0.6213   1.0000
   4.750   1.0473   0.04014   0.02964  -0.1542   0.6176   1.0000
   5.000   1.0346   0.04269   0.03222  -0.1493   0.6047   1.0000
   5.250   1.0713   0.04233   0.03177  -0.1499   0.6002   1.0000
   5.750   1.0989   0.04434   0.03375  -0.1461   0.5826   1.0000
   6.000   1.1457   0.04314   0.03247  -0.1475   0.5791   1.0000
   6.250   1.1279   0.04623   0.03564  -0.1426   0.5649   1.0000
   6.500   1.1720   0.04502   0.03437  -0.1434   0.5611   1.0000
   7.000   1.1966   0.04715   0.03656  -0.1393   0.5430   1.0000
   7.500   1.2191   0.04956   0.03904  -0.1355   0.5249   1.0000
   8.000   1.2412   0.05208   0.04165  -0.1317   0.5065   1.0000
   8.500   1.2618   0.05484   0.04452  -0.1282   0.4880   1.0000
   8.750   1.3089   0.05281   0.04249  -0.1283   0.4849   1.0000
   9.000   1.2822   0.05767   0.04748  -0.1249   0.4693   1.0000
   9.500   1.3011   0.06073   0.05066  -0.1217   0.4505   1.0000
  10.000   1.3188   0.06398   0.05406  -0.1186   0.4319   1.0000
  10.500   1.3341   0.06758   0.05781  -0.1157   0.4134   1.0000
  11.000   1.3468   0.07158   0.06197  -0.1130   0.3952   1.0000
  11.500   1.3555   0.07621   0.06677  -0.1105   0.3772   1.0000
  12.000   1.3675   0.08023   0.07094  -0.1080   0.3592   1.0000
  12.250   1.3319   0.08837   0.07919  -0.1079   0.3446   1.0000
  12.500   1.3185   0.09335   0.08425  -0.1076   0.3326   1.0000
  12.750   1.3677   0.08793   0.07886  -0.1047   0.3259   1.0000
  13.000   1.3485   0.09364   0.08465  -0.1046   0.3126   1.0000
  13.250   1.3968   0.08829   0.07925  -0.1018   0.3044   1.0000
  13.500   1.3828   0.09327   0.08434  -0.1016   0.2925   1.0000
  13.750   1.3730   0.09785   0.08904  -0.1016   0.2821   1.0000
  14.000   1.3987   0.09647   0.08767  -0.0998   0.2741   1.0000
  14.250   1.3784   0.10294   0.09428  -0.1007   0.2629   1.0000
  14.500   1.4207   0.09860   0.08987  -0.0978   0.2553   1.0000
  14.750   1.3908   0.10676   0.09822  -0.0995   0.2433   1.0000
  15.000   1.3955   0.10896   0.10049  -0.0991   0.2333   1.0000
  15.250   1.4141   0.10857   0.10007  -0.0977   0.2231   1.0000
  15.500   1.3957   0.11502   0.10667  -0.0993   0.2111   1.0000
  15.750   1.3972   0.11786   0.10954  -0.0994   0.2000   1.0000
  16.000   1.4137   0.11779   0.10937  -0.0982   0.1893   1.0000
  16.250   1.4001   0.12365   0.11537  -0.1000   0.1777   1.0000
  16.500   1.3933   0.12831   0.12008  -0.1013   0.1671   1.0000
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