RAE 101 AIRFOIL (rae101-il) Xfoil prediction polar at RE=50,000 Ncrit=9
| Details | Polar file |
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Airfoil: RAE 101 AIRFOIL (rae101-il) Reynolds number: 50,000 Max Cl/Cd: 27.92 at α=5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-rae101-il-50000.txt Download as CSV file: xf-rae101-il-50000.csv |
XFOIL Version 6.96
Calculated polar for: RAE 101 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
------ -------- --------- --------- -------- -------- --------
-9.000 -0.6598 0.06742 0.06094 -0.0208 1.0000 0.1391
-8.750 -0.6702 0.06187 0.05537 -0.0209 1.0000 0.1352
-8.500 -0.7892 0.06418 0.05666 -0.0197 1.0000 0.1264
-8.250 -0.7793 0.05923 0.05161 -0.0192 1.0000 0.1242
-8.000 -0.7760 0.05444 0.04652 -0.0185 1.0000 0.1215
-7.750 -0.7724 0.04982 0.04142 -0.0173 1.0000 0.1193
-7.500 -0.7633 0.04586 0.03695 -0.0160 1.0000 0.1195
-7.250 -0.7510 0.04253 0.03311 -0.0145 1.0000 0.1227
-7.000 -0.7366 0.03942 0.02932 -0.0127 1.0000 0.1259
-6.750 -0.7157 0.03610 0.02578 -0.0117 1.0000 0.1296
-6.500 -0.6938 0.03363 0.02304 -0.0105 1.0000 0.1377
-6.250 -0.6708 0.03122 0.02052 -0.0095 1.0000 0.1491
-6.000 -0.6457 0.02888 0.01805 -0.0085 1.0000 0.1626
-5.750 -0.6220 0.02682 0.01601 -0.0073 1.0000 0.1854
-5.500 -0.5994 0.02470 0.01419 -0.0059 1.0000 0.2199
-5.250 -0.5826 0.02244 0.01261 -0.0037 1.0000 0.2878
-5.000 -0.5792 0.02075 0.01227 0.0018 1.0000 0.4540
-4.750 -0.5721 0.02097 0.01294 0.0085 1.0000 0.5690
-4.500 -0.5587 0.02145 0.01347 0.0142 1.0000 0.6336
-4.250 -0.5417 0.02199 0.01402 0.0196 1.0000 0.6825
-4.000 -0.5230 0.02244 0.01442 0.0246 1.0000 0.7252
-3.750 -0.5033 0.02272 0.01459 0.0289 1.0000 0.7644
-3.500 -0.4687 0.02329 0.01495 0.0314 1.0000 0.7998
-3.250 -0.4263 0.02359 0.01499 0.0313 1.0000 0.8326
-3.000 -0.3716 0.02371 0.01480 0.0279 1.0000 0.8614
-2.750 -0.3106 0.02356 0.01431 0.0224 1.0000 0.8870
-2.500 -0.2555 0.02313 0.01363 0.0169 1.0000 0.9113
-2.250 -0.1874 0.02252 0.01274 0.0087 1.0000 0.9326
-2.000 -0.1337 0.02180 0.01186 0.0025 1.0000 0.9551
-1.750 -0.0696 0.02084 0.01075 -0.0059 1.0000 0.9755
-1.500 -0.0055 0.01978 0.00957 -0.0148 1.0000 0.9968
-1.250 0.0126 0.01913 0.00894 -0.0155 1.0000 1.0000
-1.000 0.0201 0.01863 0.00848 -0.0141 1.0000 1.0000
-0.750 0.0244 0.01820 0.00811 -0.0121 1.0000 1.0000
-0.500 0.0239 0.01786 0.00784 -0.0092 1.0000 1.0000
-0.250 0.0165 0.01763 0.00769 -0.0053 1.0000 1.0000
0.000 0.0000 0.01754 0.00764 0.0000 1.0000 1.0000
0.250 -0.0165 0.01763 0.00769 0.0053 1.0000 1.0000
0.500 -0.0239 0.01785 0.00784 0.0092 1.0000 1.0000
0.750 -0.0244 0.01820 0.00811 0.0121 1.0000 1.0000
1.000 -0.0200 0.01863 0.00848 0.0141 1.0000 1.0000
1.250 -0.0126 0.01913 0.00894 0.0155 1.0000 1.0000
1.500 0.0052 0.01977 0.00956 0.0149 0.9969 1.0000
1.750 0.0694 0.02084 0.01074 0.0060 0.9756 1.0000
2.000 0.1337 0.02180 0.01186 -0.0025 0.9551 1.0000
2.250 0.1875 0.02251 0.01274 -0.0087 0.9326 1.0000
2.500 0.2558 0.02313 0.01362 -0.0170 0.9114 1.0000
2.750 0.3106 0.02355 0.01430 -0.0224 0.8871 1.0000
3.000 0.3716 0.02370 0.01479 -0.0280 0.8615 1.0000
3.250 0.4263 0.02359 0.01499 -0.0313 0.8327 1.0000
3.500 0.4687 0.02329 0.01494 -0.0314 0.7999 1.0000
3.750 0.5032 0.02272 0.01459 -0.0289 0.7645 1.0000
4.000 0.5229 0.02244 0.01442 -0.0246 0.7253 1.0000
4.250 0.5416 0.02198 0.01402 -0.0196 0.6825 1.0000
4.500 0.5586 0.02144 0.01346 -0.0142 0.6336 1.0000
4.750 0.5721 0.02097 0.01294 -0.0085 0.5691 1.0000
5.000 0.5791 0.02074 0.01227 -0.0018 0.4544 1.0000
5.250 0.5826 0.02244 0.01261 0.0037 0.2878 1.0000
5.500 0.5994 0.02470 0.01419 0.0059 0.2197 1.0000
5.750 0.6219 0.02681 0.01601 0.0073 0.1854 1.0000
6.000 0.6457 0.02888 0.01805 0.0085 0.1626 1.0000
6.250 0.6708 0.03123 0.02053 0.0095 0.1491 1.0000
6.500 0.6939 0.03363 0.02304 0.0105 0.1379 1.0000
6.750 0.7157 0.03610 0.02578 0.0117 0.1297 1.0000
7.000 0.7367 0.03941 0.02931 0.0127 0.1259 1.0000
7.250 0.7511 0.04252 0.03310 0.0145 0.1228 1.0000
7.500 0.7634 0.04585 0.03693 0.0160 0.1194 1.0000
7.750 0.7725 0.04983 0.04143 0.0173 0.1193 1.0000
8.000 0.7763 0.05442 0.04650 0.0184 0.1216 1.0000
8.250 0.7798 0.05919 0.05157 0.0192 0.1242 1.0000
8.500 0.7897 0.06420 0.05667 0.0196 0.1264 1.0000
8.750 0.6700 0.06178 0.05528 0.0209 0.1353 1.0000
9.000 0.6595 0.06733 0.06085 0.0207 0.1391 1.0000
9.250 0.5962 0.10714 0.10001 -0.0199 0.3408 1.0000
9.500 0.5795 0.10967 0.10242 -0.0211 0.3327 1.0000
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Polar data table (+)
Polar graphs
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