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OAF102 AIRFOIL (oaf102-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: OAF102 AIRFOIL (oaf102-il)
Reynolds number: 50,000
Max Cl/Cd: 40.68 at α=6.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-oaf102-il-50000.txt
Download as CSV file: xf-oaf102-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: OAF102 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 =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.500  -0.4243   0.10953   0.10257  -0.0158   1.0000   0.2608
  -8.250  -0.3968   0.10352   0.09648  -0.0146   1.0000   0.2706
  -8.000  -0.4043   0.10141   0.09449  -0.0163   1.0000   0.2776
  -7.750  -0.3954   0.09825   0.09134  -0.0163   1.0000   0.2895
  -7.500  -0.3863   0.09424   0.08736  -0.0165   1.0000   0.2952
  -7.250  -0.3887   0.09205   0.08526  -0.0169   1.0000   0.3045
  -7.000  -0.3802   0.08829   0.08154  -0.0167   1.0000   0.3114
  -6.750  -0.4278   0.06840   0.06179  -0.0540   1.0000   0.1714
  -6.500  -0.4189   0.06741   0.06088  -0.0482   1.0000   0.1748
  -6.250  -0.4091   0.05757   0.05086  -0.0603   1.0000   0.1671
  -6.000  -0.3787   0.04581   0.03840  -0.0767   1.0000   0.1695
  -5.750  -0.3557   0.04196   0.03433  -0.0795   1.0000   0.1768
  -5.500  -0.3286   0.03847   0.03046  -0.0830   1.0000   0.1910
  -5.250  -0.3013   0.03567   0.02735  -0.0853   1.0000   0.2066
  -5.000  -0.2745   0.03377   0.02516  -0.0870   1.0000   0.2287
  -4.750  -0.2530   0.03284   0.02421  -0.0866   1.0000   0.2499
  -4.500  -0.2275   0.03178   0.02294  -0.0872   1.0000   0.2772
  -4.250  -0.2113   0.03189   0.02318  -0.0850   1.0000   0.2994
  -4.000  -0.1892   0.03145   0.02267  -0.0845   1.0000   0.3260
  -3.750  -0.1656   0.03101   0.02210  -0.0845   1.0000   0.3538
  -3.500  -0.1467   0.03094   0.02206  -0.0831   1.0000   0.3767
  -3.250  -0.1232   0.03059   0.02160  -0.0831   1.0000   0.4023
  -3.000  -0.0962   0.03017   0.02100  -0.0840   1.0000   0.4294
  -2.750  -0.0758   0.03001   0.02087  -0.0831   1.0000   0.4498
  -2.500  -0.0487   0.02969   0.02037  -0.0842   1.0000   0.4755
  -2.250  -0.0269   0.02951   0.02020  -0.0837   1.0000   0.4957
  -2.000   0.0003   0.02932   0.01987  -0.0849   1.0000   0.5208
  -1.750   0.0226   0.02920   0.01978  -0.0847   1.0000   0.5417
  -1.500   0.0494   0.02912   0.01960  -0.0858   1.0000   0.5668
  -1.250   0.0721   0.02910   0.01960  -0.0858   1.0000   0.5895
  -1.000   0.0974   0.02913   0.01960  -0.0866   1.0000   0.6150
  -0.750   0.1227   0.02924   0.01967  -0.0873   1.0000   0.6414
  -0.500   0.1460   0.02938   0.01982  -0.0876   1.0000   0.6672
  -0.250   0.1697   0.02957   0.02005  -0.0880   1.0000   0.6957
   0.000   0.1928   0.02981   0.02034  -0.0883   1.0000   0.7274
   0.250   0.2151   0.03007   0.02070  -0.0884   1.0000   0.7647
   0.500   0.2396   0.03015   0.02098  -0.0885   0.9967   0.8139
   0.750   0.2583   0.02973   0.02088  -0.0875   0.9874   0.9371
   1.000   0.3169   0.03074   0.02160  -0.0973   0.9772   1.0000
   1.250   0.3712   0.03186   0.02244  -0.1048   0.9683   1.0000
   1.500   0.4135   0.03290   0.02331  -0.1094   0.9587   1.0000
   1.750   0.4480   0.03396   0.02427  -0.1124   0.9488   1.0000
   2.000   0.4892   0.03506   0.02530  -0.1162   0.9396   1.0000
   2.250   0.5219   0.03615   0.02637  -0.1186   0.9296   1.0000
   2.500   0.5504   0.03730   0.02753  -0.1202   0.9193   1.0000
   2.750   0.5878   0.03846   0.02871  -0.1231   0.9098   1.0000
   3.000   0.6163   0.03963   0.02994  -0.1245   0.8991   1.0000
   3.250   0.6402   0.04091   0.03130  -0.1252   0.8882   1.0000
   3.500   0.6720   0.04215   0.03263  -0.1270   0.8777   1.0000
   3.750   0.7077   0.04331   0.03391  -0.1293   0.8667   1.0000
   4.250   0.7509   0.04606   0.03693  -0.1296   0.8416   1.0000
   4.500   0.7779   0.04736   0.03838  -0.1304   0.8282   1.0000
   4.750   0.8061   0.04857   0.03977  -0.1310   0.8137   1.0000
   5.000   0.8354   0.04963   0.04108  -0.1315   0.7974   1.0000
   5.250   0.8846   0.04949   0.04125  -0.1331   0.7763   1.0000
   5.500   1.0131   0.03792   0.03035  -0.1292   0.7116   1.0000
   5.750   1.0585   0.03295   0.02571  -0.1227   0.6721   1.0000
   6.000   1.0907   0.02779   0.02078  -0.1137   0.6075   1.0000
   6.250   1.0829   0.02662   0.01796  -0.1012   0.3239   1.0000
   6.500   1.0763   0.03066   0.02051  -0.0966   0.2229   1.0000
   6.750   1.0884   0.03331   0.02273  -0.0940   0.1849   1.0000
   7.000   1.1153   0.03560   0.02472  -0.0928   0.1613   1.0000
   7.250   1.1491   0.03785   0.02688  -0.0924   0.1433   1.0000
   7.500   1.1867   0.04061   0.02956  -0.0927   0.1327   1.0000
   7.750   1.2178   0.04340   0.03236  -0.0926   0.1236   1.0000
   8.000   1.2440   0.04657   0.03592  -0.0918   0.1187   1.0000
   8.250   1.2678   0.04999   0.03985  -0.0906   0.1162   1.0000
   8.500   1.2876   0.05349   0.04375  -0.0894   0.1133   1.0000
   8.750   1.3093   0.05733   0.04767  -0.0887   0.1095   1.0000
   9.000   1.3251   0.06197   0.05266  -0.0875   0.1086   1.0000
   9.250   1.3244   0.06588   0.05746  -0.0843   0.1106   1.0000
   9.500   1.3179   0.07095   0.06325  -0.0813   0.1135   1.0000
   9.750   1.3081   0.07622   0.06903  -0.0788   0.1163   1.0000
  10.000   1.2986   0.08150   0.07465  -0.0768   0.1186   1.0000
  10.250   1.2931   0.08703   0.08039  -0.0753   0.1205   1.0000
  10.500   1.2706   0.09195   0.08565  -0.0733   0.1237   1.0000
  10.750   1.2022   0.09779   0.09183  -0.0722   0.1268   1.0000
  11.000   1.1540   0.10648   0.10067  -0.0760   0.1297   1.0000
  11.250   1.1375   0.11464   0.10886  -0.0790   0.1342   1.0000
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