UNIVERSITY OF ILLINOIS UI-1720 AIRFOIL (ui1720-il) Xfoil prediction polar at RE=50,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: UNIVERSITY OF ILLINOIS UI-1720 AIRFOIL (ui1720-il) Reynolds number: 50,000 Max Cl/Cd: 17.85 at α=2° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-ui1720-il-50000-n5.txt Download as CSV file: xf-ui1720-il-50000-n5.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 = 0.050 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-11.250 -0.3507 0.14322 0.13665 0.0053 1.0004 0.1039
-11.000 -0.3537 0.14127 0.13478 0.0027 1.0004 0.1063
-10.750 -0.3575 0.13914 0.13274 0.0001 1.0004 0.1069
-10.500 -0.3320 0.13348 0.12719 0.0022 1.0004 0.1110
-10.250 -0.3252 0.13045 0.12425 0.0013 1.0004 0.1145
-10.000 -0.3242 0.12781 0.12172 -0.0006 1.0004 0.1173
-9.750 -0.3265 0.12534 0.11936 -0.0030 1.0004 0.1185
-9.250 -0.2889 0.11769 0.11035 -0.0067 0.5267 0.1246
-9.000 -0.3008 0.11131 0.10401 -0.0133 0.5338 0.0957
-8.500 -0.2816 0.10442 0.09662 -0.0149 0.4708 0.0936
-8.250 -0.2766 0.10087 0.09299 -0.0169 0.4568 0.0935
-8.000 -0.2712 0.09727 0.08929 -0.0188 0.4458 0.0928
-7.750 -0.2670 0.09361 0.08559 -0.0211 0.4369 0.0930
-7.500 -0.2650 0.08993 0.08186 -0.0238 0.4305 0.0939
-7.250 -0.2617 0.08588 0.07778 -0.0275 0.4248 0.0944
-7.000 -0.2554 0.08122 0.07310 -0.0325 0.4193 0.0948
-6.750 -0.2491 0.07458 0.06637 -0.0419 0.4161 0.0959
-6.500 -0.2318 0.07361 0.06529 -0.0398 0.4081 0.0987
-6.250 -0.2171 0.06935 0.06096 -0.0444 0.4031 0.1005
-6.000 -0.2008 0.06409 0.05554 -0.0509 0.3990 0.1029
-5.750 -0.1821 0.05707 0.04817 -0.0599 0.3958 0.1051
-5.500 -0.1599 0.05188 0.04256 -0.0657 0.3920 0.1081
-5.250 -0.1362 0.04991 0.04041 -0.0669 0.3876 0.1109
-5.000 -0.1097 0.04663 0.03679 -0.0702 0.3834 0.1146
-4.750 -0.0803 0.04250 0.03198 -0.0746 0.3797 0.1181
-4.500 -0.0527 0.04064 0.02986 -0.0758 0.3758 0.1212
-4.250 -0.0247 0.03920 0.02818 -0.0766 0.3723 0.1240
-4.000 0.0051 0.03765 0.02619 -0.0779 0.3692 0.1279
-3.750 0.0361 0.03592 0.02401 -0.0793 0.3658 0.1317
-3.500 0.0655 0.03473 0.02259 -0.0800 0.3623 0.1344
-3.250 0.0938 0.03404 0.02182 -0.0801 0.3589 0.1384
-3.000 0.1227 0.03329 0.02082 -0.0804 0.3560 0.1426
-2.750 0.1521 0.03254 0.01970 -0.0806 0.3534 0.1465
-2.500 0.1803 0.03207 0.01909 -0.0806 0.3513 0.1510
-2.250 0.2084 0.03179 0.01866 -0.0805 0.3496 0.1567
-2.000 0.2370 0.03147 0.01820 -0.0804 0.3479 0.1626
-1.750 0.2644 0.03127 0.01806 -0.0802 0.3462 0.1687
-1.500 0.2919 0.03121 0.01796 -0.0799 0.3446 0.1777
-1.250 0.3196 0.03118 0.01794 -0.0797 0.3430 0.1870
-1.000 0.3474 0.03124 0.01796 -0.0795 0.3416 0.2003
-0.750 0.3748 0.03134 0.01811 -0.0794 0.3402 0.2172
-0.500 0.4020 0.03146 0.01836 -0.0793 0.3389 0.2402
-0.250 0.4289 0.03162 0.01868 -0.0791 0.3377 0.2748
0.000 0.4554 0.03176 0.01910 -0.0790 0.3365 0.3275
0.250 0.4809 0.03188 0.01959 -0.0786 0.3352 0.4104
0.500 0.5040 0.03192 0.02012 -0.0776 0.3340 0.5304
0.750 0.5211 0.03174 0.02056 -0.0748 0.3328 0.7078
1.000 0.5489 0.03195 0.02087 -0.0741 0.3315 0.9996
1.250 0.5758 0.03282 0.02146 -0.0741 0.3304 0.9996
1.500 0.6004 0.03389 0.02248 -0.0741 0.3292 0.9996
1.750 0.6246 0.03505 0.02360 -0.0740 0.3283 0.9996
2.000 0.6482 0.03631 0.02483 -0.0739 0.3275 0.9996
2.250 0.6710 0.03767 0.02618 -0.0738 0.3268 0.9996
2.500 0.6930 0.03914 0.02766 -0.0737 0.3260 0.9996
2.750 0.7140 0.04074 0.02928 -0.0737 0.3254 0.9996
3.000 0.7340 0.04248 0.03107 -0.0736 0.3248 0.9996
3.250 0.7527 0.04438 0.03302 -0.0736 0.3240 0.9996
3.500 0.7700 0.04642 0.03512 -0.0737 0.3228 0.9996
3.750 0.7860 0.04857 0.03733 -0.0737 0.3215 0.9996
4.000 0.8007 0.05087 0.03969 -0.0739 0.3202 0.9996
4.250 0.8131 0.05353 0.04241 -0.0741 0.3193 0.9996
4.500 0.8214 0.05673 0.04571 -0.0747 0.3187 0.9996
4.750 0.8254 0.06049 0.04959 -0.0756 0.3183 0.9996
5.000 0.8277 0.06434 0.05351 -0.0766 0.3174 0.9996
5.250 0.8325 0.06772 0.05693 -0.0774 0.3161 0.9996
5.500 0.8399 0.07071 0.05992 -0.0779 0.3148 0.9996
5.750 0.8440 0.07420 0.06343 -0.0787 0.3138 0.9996
6.000 0.8461 0.07779 0.06703 -0.0794 0.3131 0.9996
6.250 0.8625 0.07974 0.06894 -0.0788 0.3120 0.9996
6.500 0.8185 0.08893 0.07825 -0.0848 0.3097 0.9996
6.750 0.8043 0.09424 0.08361 -0.0873 0.3083 0.9996
7.000 0.7962 0.09867 0.08807 -0.0891 0.3062 0.9996
7.250 0.7939 0.10243 0.09183 -0.0902 0.3041 0.9996
7.500 0.7950 0.10596 0.09539 -0.0911 0.3028 0.9996
7.750 0.7972 0.10939 0.09882 -0.0919 0.3017 0.9996
|
Polar data table (+)
Polar graphs
<< Back to UNIVERSITY OF ILLINOIS UI-1720 AIRFOIL (ui1720-il)