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OAF102 AIRFOIL (oaf102-il) Xfoil prediction polar at RE=100,000 Ncrit=9


Details Polar file
Airfoil: OAF102 AIRFOIL (oaf102-il)
Reynolds number: 100,000
Max Cl/Cd: 61.91 at α=5.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-oaf102-il-100000.txt
Download as CSV file: xf-oaf102-il-100000.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 =     0.100 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.500  -0.4560   0.09879   0.09401  -0.0293   1.0000   0.1427
  -8.250  -0.4253   0.09427   0.08943  -0.0248   1.0000   0.1470
  -8.000  -0.5030   0.05444   0.04947  -0.0731   1.0000   0.0735
  -7.750  -0.4918   0.04493   0.03941  -0.0838   1.0000   0.0731
  -7.500  -0.4667   0.03737   0.03086  -0.0925   1.0000   0.0747
  -7.250  -0.4447   0.03337   0.02665  -0.0945   1.0000   0.0770
  -7.000  -0.4217   0.03091   0.02396  -0.0955   1.0000   0.0803
  -6.750  -0.3958   0.02852   0.02099  -0.0971   1.0000   0.0861
  -6.500  -0.3736   0.02633   0.01862  -0.0975   1.0000   0.0910
  -6.250  -0.3512   0.02503   0.01709  -0.0975   1.0000   0.0979
  -6.000  -0.3290   0.02373   0.01572  -0.0975   1.0000   0.1062
  -5.750  -0.3045   0.02252   0.01428  -0.0978   1.0000   0.1166
  -5.500  -0.2807   0.02173   0.01344  -0.0980   1.0000   0.1305
  -5.250  -0.2567   0.02111   0.01287  -0.0982   1.0000   0.1470
  -5.000  -0.2318   0.02075   0.01248  -0.0986   1.0000   0.1694
  -4.750  -0.2066   0.02050   0.01233  -0.0990   1.0000   0.1946
  -4.500  -0.1812   0.02040   0.01226  -0.0995   1.0000   0.2218
  -4.250  -0.1554   0.02043   0.01228  -0.1000   0.9997   0.2487
  -4.000  -0.1118   0.02039   0.01228  -0.1036   0.9944   0.2804
  -3.750  -0.0703   0.02038   0.01224  -0.1067   0.9881   0.3091
  -3.500  -0.0255   0.02037   0.01218  -0.1104   0.9832   0.3372
  -3.250   0.0124   0.02037   0.01226  -0.1127   0.9759   0.3586
  -3.000   0.0576   0.02027   0.01205  -0.1165   0.9711   0.3835
  -2.750   0.0948   0.02021   0.01204  -0.1186   0.9637   0.4027
  -2.500   0.1386   0.02007   0.01189  -0.1219   0.9585   0.4249
  -2.250   0.1783   0.01996   0.01175  -0.1245   0.9518   0.4464
  -2.000   0.2199   0.01978   0.01162  -0.1273   0.9458   0.4674
  -1.750   0.2658   0.01953   0.01135  -0.1310   0.9417   0.4917
  -1.500   0.2964   0.01953   0.01141  -0.1317   0.9324   0.5107
  -1.250   0.3380   0.01931   0.01121  -0.1344   0.9275   0.5343
  -1.000   0.3674   0.01933   0.01129  -0.1349   0.9184   0.5536
  -0.750   0.4049   0.01914   0.01117  -0.1366   0.9128   0.5749
  -0.500   0.4334   0.01924   0.01131  -0.1370   0.9038   0.5966
  -0.250   0.4673   0.01908   0.01127  -0.1379   0.8978   0.6188
   0.000   0.4944   0.01920   0.01147  -0.1379   0.8886   0.6425
   0.250   0.5258   0.01906   0.01147  -0.1382   0.8824   0.6690
   0.500   0.5506   0.01920   0.01175  -0.1376   0.8730   0.6983
   0.750   0.5793   0.01902   0.01174  -0.1371   0.8667   0.7364
   1.000   0.5997   0.01906   0.01205  -0.1355   0.8567   0.7876
   1.250   0.6149   0.01860   0.01192  -0.1319   0.8492   1.0000
   1.500   0.6487   0.01898   0.01218  -0.1338   0.8399   1.0000
   1.750   0.6797   0.01927   0.01240  -0.1345   0.8323   1.0000
   2.000   0.7088   0.01959   0.01269  -0.1348   0.8238   1.0000
   2.250   0.7368   0.01998   0.01309  -0.1349   0.8153   1.0000
   2.500   0.7657   0.02021   0.01334  -0.1347   0.8075   1.0000
   2.750   0.7924   0.02068   0.01386  -0.1347   0.7981   1.0000
   3.000   0.8214   0.02081   0.01404  -0.1343   0.7911   1.0000
   3.250   0.8473   0.02134   0.01465  -0.1341   0.7807   1.0000
   3.500   0.8761   0.02143   0.01480  -0.1335   0.7738   1.0000
   3.750   0.9021   0.02173   0.01522  -0.1329   0.7623   1.0000
   4.000   0.9278   0.02097   0.01447  -0.1300   0.7457   1.0000
   4.250   0.9504   0.01983   0.01330  -0.1260   0.7192   1.0000
   4.500   0.9745   0.01893   0.01236  -0.1228   0.6949   1.0000
   4.750   0.9996   0.01849   0.01197  -0.1208   0.6724   1.0000
   5.000   1.0246   0.01809   0.01163  -0.1189   0.6466   1.0000
   5.250   1.0489   0.01766   0.01125  -0.1168   0.6125   1.0000
   5.500   1.0711   0.01730   0.01085  -0.1143   0.5496   1.0000
   5.750   1.0698   0.02002   0.01143  -0.1091   0.2189   1.0000
   6.000   1.0767   0.02314   0.01365  -0.1065   0.1339   1.0000
   6.250   1.0923   0.02489   0.01525  -0.1046   0.1122   1.0000
   6.500   1.1091   0.02657   0.01685  -0.1027   0.1002   1.0000
   6.750   1.1280   0.02836   0.01851  -0.1012   0.0913   1.0000
   7.000   1.1500   0.02996   0.02005  -0.1002   0.0835   1.0000
   7.250   1.1766   0.03215   0.02221  -0.0995   0.0787   1.0000
   7.500   1.2023   0.03384   0.02403  -0.0988   0.0735   1.0000
   7.750   1.2297   0.03637   0.02643  -0.0988   0.0690   1.0000
   8.000   1.2572   0.03927   0.02959  -0.0983   0.0671   1.0000
   8.250   1.2803   0.04167   0.03241  -0.0970   0.0653   1.0000
   8.500   1.3007   0.04417   0.03530  -0.0957   0.0628   1.0000
   8.750   1.3196   0.04721   0.03871  -0.0942   0.0615   1.0000
   9.000   1.3352   0.05104   0.04302  -0.0924   0.0617   1.0000
   9.250   1.3458   0.05545   0.04796  -0.0901   0.0629   1.0000
   9.500   1.3526   0.06010   0.05306  -0.0877   0.0642   1.0000
   9.750   1.3554   0.06472   0.05811  -0.0853   0.0650   1.0000
  10.000   1.3546   0.06946   0.06320  -0.0828   0.0658   1.0000
  10.250   1.3535   0.07482   0.06880  -0.0809   0.0668   1.0000
  10.500   1.2272   0.07093   0.06568  -0.0619   0.0710   1.0000
  10.750   1.1683   0.07674   0.07192  -0.0566   0.0751   1.0000
  11.000   1.1348   0.08230   0.07768  -0.0548   0.0767   1.0000
  11.250   1.1067   0.08821   0.08374  -0.0541   0.0783   1.0000
  11.500   1.0545   0.09622   0.09205  -0.0553   0.0843   1.0000
  11.750   1.0062   0.10407   0.10004  -0.0582   0.0844   1.0000
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