RAE 101 AIRFOIL (rae101-il) Xfoil prediction polar at RE=50,000 Ncrit=5
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Airfoil: RAE 101 AIRFOIL (rae101-il) Reynolds number: 50,000 Max Cl/Cd: 27.44 at α=4.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-rae101-il-50000-n5.txt Download as CSV file: xf-rae101-il-50000-n5.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 = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-11.000 -0.7123 0.09231 0.08506 -0.0126 1.0000 0.0471
-10.750 -0.7287 0.08522 0.07795 -0.0175 1.0000 0.0468
-10.500 -0.7464 0.07896 0.07165 -0.0216 1.0000 0.0465
-10.250 -0.7651 0.07363 0.06623 -0.0244 1.0000 0.0462
-10.000 -0.7844 0.06915 0.06163 -0.0253 1.0000 0.0460
-9.750 -0.8015 0.06523 0.05753 -0.0247 1.0000 0.0460
-9.500 -0.8139 0.06119 0.05320 -0.0239 1.0000 0.0463
-9.250 -0.8221 0.05723 0.04885 -0.0227 1.0000 0.0469
-9.000 -0.8258 0.05353 0.04461 -0.0212 1.0000 0.0478
-8.750 -0.8176 0.05024 0.04124 -0.0204 1.0000 0.0494
-8.500 -0.8076 0.04733 0.03808 -0.0193 1.0000 0.0510
-8.250 -0.7954 0.04432 0.03474 -0.0181 1.0000 0.0523
-8.000 -0.7807 0.04143 0.03148 -0.0169 1.0000 0.0541
-7.750 -0.7641 0.03891 0.02849 -0.0156 1.0000 0.0577
-7.500 -0.7451 0.03652 0.02582 -0.0146 1.0000 0.0614
-7.250 -0.7239 0.03450 0.02365 -0.0137 1.0000 0.0651
-7.000 -0.7012 0.03260 0.02147 -0.0127 1.0000 0.0703
-6.750 -0.6802 0.03094 0.01978 -0.0117 1.0000 0.0770
-6.500 -0.6581 0.02941 0.01804 -0.0105 1.0000 0.0846
-6.250 -0.6390 0.02795 0.01662 -0.0092 1.0000 0.0946
-6.000 -0.6203 0.02655 0.01518 -0.0078 1.0000 0.1077
-5.750 -0.6022 0.02517 0.01379 -0.0064 1.0000 0.1263
-5.500 -0.5858 0.02365 0.01248 -0.0050 1.0000 0.1556
-5.250 -0.5708 0.02194 0.01123 -0.0035 1.0000 0.2129
-5.000 -0.5582 0.02044 0.01053 -0.0015 1.0000 0.3270
-4.750 -0.5428 0.01978 0.01023 0.0010 1.0000 0.4287
-4.500 -0.5258 0.01943 0.00999 0.0035 1.0000 0.4954
-4.250 -0.5082 0.01920 0.00983 0.0060 1.0000 0.5475
-4.000 -0.4901 0.01903 0.00970 0.0086 1.0000 0.5912
-3.750 -0.4715 0.01888 0.00955 0.0111 1.0000 0.6291
-3.500 -0.4523 0.01872 0.00935 0.0135 1.0000 0.6624
-3.250 -0.4323 0.01855 0.00916 0.0157 1.0000 0.6917
-3.000 -0.4115 0.01837 0.00894 0.0176 1.0000 0.7180
-2.750 -0.3897 0.01818 0.00873 0.0193 1.0000 0.7418
-2.500 -0.3682 0.01797 0.00845 0.0208 1.0000 0.7654
-2.250 -0.3445 0.01781 0.00825 0.0219 1.0000 0.7873
-2.000 -0.3207 0.01764 0.00804 0.0229 1.0000 0.8103
-1.750 -0.2938 0.01753 0.00790 0.0232 1.0000 0.8327
-1.500 -0.2644 0.01745 0.00776 0.0229 1.0000 0.8561
-1.250 -0.2315 0.01742 0.00769 0.0218 1.0000 0.8805
-1.000 -0.1919 0.01744 0.00767 0.0192 1.0000 0.9035
-0.750 -0.1507 0.01747 0.00764 0.0161 1.0000 0.9276
-0.500 -0.1036 0.01751 0.00763 0.0116 1.0000 0.9494
-0.250 -0.0564 0.01753 0.00763 0.0068 1.0000 0.9721
0.000 0.0000 0.01751 0.00761 0.0000 0.9901 0.9901
0.250 0.0564 0.01753 0.00763 -0.0068 0.9721 1.0000
0.500 0.1035 0.01751 0.00763 -0.0116 0.9495 1.0000
0.750 0.1505 0.01747 0.00764 -0.0161 0.9277 1.0000
1.000 0.1917 0.01744 0.00767 -0.0192 0.9036 1.0000
1.250 0.2314 0.01742 0.00769 -0.0218 0.8805 1.0000
1.500 0.2643 0.01745 0.00776 -0.0229 0.8562 1.0000
1.750 0.2937 0.01753 0.00790 -0.0232 0.8327 1.0000
2.000 0.3206 0.01764 0.00804 -0.0229 0.8103 1.0000
2.250 0.3444 0.01780 0.00824 -0.0219 0.7874 1.0000
2.500 0.3681 0.01797 0.00844 -0.0208 0.7655 1.0000
2.750 0.3897 0.01818 0.00873 -0.0193 0.7418 1.0000
3.000 0.4115 0.01837 0.00894 -0.0176 0.7181 1.0000
3.250 0.4322 0.01855 0.00916 -0.0157 0.6917 1.0000
3.500 0.4522 0.01872 0.00935 -0.0135 0.6624 1.0000
3.750 0.4714 0.01888 0.00955 -0.0111 0.6292 1.0000
4.000 0.4900 0.01903 0.00969 -0.0086 0.5912 1.0000
4.250 0.5081 0.01920 0.00983 -0.0060 0.5475 1.0000
4.500 0.5257 0.01943 0.00999 -0.0034 0.4955 1.0000
4.750 0.5427 0.01978 0.01022 -0.0009 0.4285 1.0000
5.000 0.5581 0.02044 0.01053 0.0015 0.3272 1.0000
5.250 0.5708 0.02194 0.01123 0.0035 0.2129 1.0000
5.500 0.5857 0.02365 0.01248 0.0050 0.1555 1.0000
5.750 0.6023 0.02517 0.01378 0.0064 0.1263 1.0000
6.000 0.6203 0.02656 0.01518 0.0078 0.1076 1.0000
6.250 0.6390 0.02795 0.01662 0.0092 0.0945 1.0000
6.500 0.6582 0.02941 0.01803 0.0105 0.0846 1.0000
6.750 0.6804 0.03094 0.01978 0.0117 0.0771 1.0000
7.000 0.7014 0.03261 0.02147 0.0127 0.0704 1.0000
7.250 0.7240 0.03450 0.02365 0.0137 0.0652 1.0000
7.500 0.7452 0.03652 0.02581 0.0146 0.0614 1.0000
7.750 0.7643 0.03892 0.02850 0.0156 0.0577 1.0000
8.000 0.7808 0.04143 0.03147 0.0168 0.0542 1.0000
8.250 0.7956 0.04430 0.03472 0.0180 0.0524 1.0000
8.500 0.8077 0.04734 0.03809 0.0192 0.0510 1.0000
8.750 0.8179 0.05023 0.04123 0.0203 0.0494 1.0000
9.000 0.8261 0.05351 0.04460 0.0212 0.0478 1.0000
9.250 0.8225 0.05723 0.04885 0.0227 0.0469 1.0000
9.500 0.8144 0.06118 0.05320 0.0238 0.0463 1.0000
9.750 0.8023 0.06521 0.05751 0.0246 0.0460 1.0000
10.000 0.7851 0.06913 0.06162 0.0252 0.0460 1.0000
10.250 0.7652 0.07369 0.06629 0.0242 0.0461 1.0000
10.500 0.7461 0.07911 0.07180 0.0214 0.0464 1.0000
10.750 0.7297 0.08521 0.07795 0.0174 0.0469 1.0000
11.000 0.7155 0.09187 0.08461 0.0130 0.0472 1.0000
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Polar data table (+)
Polar graphs
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