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UNIVERSITY OF ILLINOIS UI-1720 AIRFOIL (ui1720-il) Xfoil prediction polar at RE=200,000 Ncrit=9


Details Polar file
Airfoil: UNIVERSITY OF ILLINOIS UI-1720 AIRFOIL (ui1720-il)
Reynolds number: 200,000
Max Cl/Cd: 30.19 at α=2.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-ui1720-il-200000.txt
Download as CSV file: xf-ui1720-il-200000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: UNIVERSITY OF ILLINOIS UI-1720 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
  ------ -------- --------- --------- -------- -------- --------
  -9.250  -0.2740   0.10710   0.10136  -0.0157   0.3485   0.0654
  -9.000  -0.2686   0.10398   0.09825  -0.0173   0.3460   0.0678
  -8.750  -0.2888   0.09816   0.09252  -0.0261   0.3479   0.0700
  -8.500  -0.2620   0.09681   0.09109  -0.0212   0.3418   0.0714
  -8.250  -0.2510   0.09439   0.08865  -0.0215   0.3388   0.0738
  -8.000  -0.2791   0.08753   0.08194  -0.0352   0.3417   0.0775
  -7.750  -0.2544   0.08567   0.07998  -0.0296   0.3368   0.0785
  -7.500  -0.2368   0.08416   0.07839  -0.0276   0.3335   0.0804
  -7.000  -0.2334   0.07549   0.06973  -0.0383   0.3314   0.0875
  -6.750  -0.2166   0.07428   0.06848  -0.0360   0.3292   0.0897
  -6.500  -0.2250   0.03629   0.02908  -0.0846   0.3350   0.0778
  -6.250  -0.1901   0.04916   0.04265  -0.0762   0.3295   0.0962
  -6.000  -0.1739   0.03896   0.03212  -0.0810   0.3285   0.0839
  -5.750  -0.1508   0.03108   0.02311  -0.0864   0.3273   0.0870
  -5.500  -0.1248   0.03183   0.02419  -0.0853   0.3249   0.0896
  -5.250  -0.0982   0.02997   0.02199  -0.0860   0.3229   0.0927
  -5.000  -0.0703   0.02780   0.01915  -0.0869   0.3211   0.0956
  -4.750  -0.0440   0.02633   0.01757  -0.0870   0.3193   0.0980
  -4.500  -0.0169   0.02579   0.01698  -0.0868   0.3177   0.1004
  -4.250   0.0110   0.02501   0.01596  -0.0868   0.3162   0.1030
  -4.000   0.0393   0.02422   0.01484  -0.0868   0.3148   0.1053
  -3.750   0.0678   0.02369   0.01397  -0.0867   0.3133   0.1073
  -3.500   0.0952   0.02300   0.01321  -0.0866   0.3117   0.1092
  -3.250   0.1231   0.02255   0.01273  -0.0864   0.3108   0.1116
  -3.000   0.1513   0.02218   0.01229  -0.0861   0.3097   0.1140
  -2.750   0.1797   0.02182   0.01183  -0.0858   0.3088   0.1161
  -2.500   0.2081   0.02163   0.01150  -0.0855   0.3080   0.1184
  -2.250   0.2361   0.02120   0.01106  -0.0853   0.3071   0.1211
  -2.000   0.2639   0.02100   0.01089  -0.0850   0.3063   0.1239
  -1.750   0.2920   0.02091   0.01080  -0.0847   0.3056   0.1272
  -1.500   0.3201   0.02091   0.01076  -0.0843   0.3049   0.1313
  -1.250   0.3478   0.02077   0.01067  -0.0840   0.3042   0.1360
  -1.000   0.3755   0.02080   0.01077  -0.0837   0.3036   0.1419
  -0.750   0.4033   0.02086   0.01085  -0.0834   0.3031   0.1491
  -0.500   0.4309   0.02093   0.01102  -0.0831   0.3026   0.1601
  -0.250   0.4585   0.02102   0.01123  -0.0829   0.3021   0.1775
   0.000   0.4860   0.02109   0.01150  -0.0827   0.3017   0.2100
   0.250   0.5132   0.02115   0.01187  -0.0825   0.3013   0.2795
   0.500   0.5398   0.02124   0.01236  -0.0823   0.3010   0.3812
   0.750   0.5656   0.02139   0.01292  -0.0818   0.3007   0.4950
   1.000   0.5893   0.02148   0.01353  -0.0807   0.3005   0.6361
   1.250   0.6140   0.02125   0.01401  -0.0792   0.3003   0.9996
   1.500   0.6411   0.02188   0.01457  -0.0789   0.3000   0.9996
   1.750   0.6681   0.02248   0.01511  -0.0787   0.2993   0.9996
   2.000   0.6948   0.02311   0.01569  -0.0784   0.2988   0.9996
   2.250   0.7209   0.02388   0.01645  -0.0781   0.2986   0.9996
   2.500   0.7465   0.02473   0.01731  -0.0777   0.2985   0.9996
   2.750   0.7716   0.02567   0.01828  -0.0773   0.2985   0.9996
   3.000   0.7961   0.02672   0.01939  -0.0769   0.2985   0.9996
   3.250   0.8197   0.02793   0.02068  -0.0765   0.2988   0.9996
   3.500   0.8418   0.02940   0.02227  -0.0760   0.2992   0.9996
   3.750   0.8618   0.03131   0.02436  -0.0756   0.3002   0.9996
   4.000   0.8798   0.03364   0.02690  -0.0751   0.3015   0.9996
   4.250   0.8990   0.03553   0.02890  -0.0747   0.3015   0.9996
   4.500   0.9102   0.03921   0.03285  -0.0746   0.3043   0.9996
   4.750   0.9223   0.04261   0.03643  -0.0745   0.3066   0.9996
   5.000   0.9350   0.04574   0.03967  -0.0745   0.3080   0.9996
   5.250   0.9506   0.04849   0.04247  -0.0743   0.3093   0.9996
   5.500   0.9678   0.05106   0.04505  -0.0742   0.3102   0.9996
   5.750   0.9982   0.05187   0.04576  -0.0739   0.3127   0.9996
   6.000   0.8090   0.08252   0.07740  -0.0925   0.3158   0.9996
   6.250   0.8304   0.08393   0.07876  -0.0913   0.3147   0.9996
   6.500   0.8537   0.08530   0.08011  -0.0899   0.3139   0.9996
   6.750   0.8788   0.08671   0.08149  -0.0882   0.3132   0.9996
   7.000   0.9065   0.08809   0.08284  -0.0863   0.3127   0.9996
   7.250   0.9375   0.08941   0.08414  -0.0838   0.3122   0.9996
   7.500   0.9734   0.09089   0.08558  -0.0813   0.3118   0.9996
   7.750   0.8086   0.10601   0.10084  -0.0969   0.2867   0.9996
   8.000   0.8169   0.10827   0.10308  -0.0970   0.2826   0.9996
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