RAE 101 AIRFOIL (rae101-il) Xfoil prediction polar at RE=500,000 Ncrit=9
| Details | Polar file |
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Airfoil: RAE 101 AIRFOIL (rae101-il) Reynolds number: 500,000 Max Cl/Cd: 60.26 at α=3.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-rae101-il-500000.txt Download as CSV file: xf-rae101-il-500000.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 = 0.500 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-12.250 -0.7972 0.07888 0.07643 -0.0141 1.0000 0.0149
-12.000 -0.8240 0.06925 0.06669 -0.0216 1.0000 0.0146
-11.750 -0.8492 0.06146 0.05876 -0.0275 1.0000 0.0145
-11.500 -0.8755 0.05495 0.05207 -0.0315 1.0000 0.0143
-11.250 -0.9039 0.04939 0.04630 -0.0330 1.0000 0.0138
-11.000 -0.9302 0.04508 0.04176 -0.0316 1.0000 0.0138
-10.750 -0.9448 0.04249 0.03896 -0.0284 1.0000 0.0140
-10.500 -0.9599 0.03864 0.03479 -0.0251 1.0000 0.0139
-10.250 -0.9646 0.03510 0.03091 -0.0224 1.0000 0.0140
-10.000 -0.9629 0.03189 0.02734 -0.0200 1.0000 0.0141
-9.750 -0.9547 0.02931 0.02444 -0.0179 1.0000 0.0143
-9.500 -0.9421 0.02724 0.02209 -0.0162 1.0000 0.0147
-9.250 -0.9269 0.02541 0.02001 -0.0146 1.0000 0.0150
-9.000 -0.9089 0.02414 0.01854 -0.0133 1.0000 0.0154
-8.750 -0.8912 0.02257 0.01676 -0.0120 1.0000 0.0157
-8.500 -0.8770 0.02005 0.01408 -0.0102 1.0000 0.0166
-8.250 -0.8571 0.01916 0.01315 -0.0092 1.0000 0.0173
-8.000 -0.8360 0.01849 0.01242 -0.0082 1.0000 0.0183
-7.750 -0.8145 0.01785 0.01171 -0.0072 1.0000 0.0195
-7.500 -0.7935 0.01710 0.01087 -0.0060 1.0000 0.0204
-7.250 -0.7715 0.01654 0.01022 -0.0049 1.0000 0.0211
-7.000 -0.7572 0.01490 0.00850 -0.0027 1.0000 0.0228
-6.750 -0.7363 0.01432 0.00791 -0.0016 1.0000 0.0245
-6.500 -0.7144 0.01388 0.00742 -0.0005 1.0000 0.0265
-6.250 -0.6916 0.01354 0.00703 0.0005 1.0000 0.0281
-6.000 -0.6743 0.01253 0.00598 0.0023 1.0000 0.0316
-5.750 -0.6525 0.01211 0.00554 0.0035 1.0000 0.0348
-5.500 -0.6313 0.01163 0.00501 0.0048 1.0000 0.0393
-5.250 -0.6101 0.01118 0.00458 0.0060 1.0000 0.0457
-5.000 -0.5892 0.01071 0.00414 0.0073 1.0000 0.0552
-4.750 -0.5683 0.01029 0.00378 0.0085 1.0000 0.0709
-4.500 -0.5482 0.00980 0.00344 0.0099 1.0000 0.1010
-4.250 -0.5291 0.00925 0.00312 0.0113 1.0000 0.1514
-4.000 -0.5072 0.00856 0.00281 0.0120 0.9992 0.2328
-3.750 -0.4712 0.00782 0.00258 0.0096 0.9946 0.3520
-3.500 -0.4349 0.00747 0.00243 0.0075 0.9891 0.4141
-3.250 -0.3963 0.00724 0.00232 0.0051 0.9842 0.4555
-3.000 -0.3605 0.00702 0.00221 0.0034 0.9773 0.4884
-2.750 -0.3221 0.00684 0.00210 0.0011 0.9716 0.5180
-2.500 -0.2868 0.00668 0.00200 -0.0004 0.9627 0.5445
-2.250 -0.2505 0.00651 0.00191 -0.0022 0.9533 0.5704
-2.000 -0.2160 0.00636 0.00182 -0.0035 0.9403 0.5953
-1.750 -0.1846 0.00624 0.00174 -0.0040 0.9230 0.6193
-1.500 -0.1555 0.00614 0.00167 -0.0040 0.9021 0.6428
-1.250 -0.1284 0.00607 0.00162 -0.0036 0.8796 0.6651
-1.000 -0.1024 0.00603 0.00157 -0.0029 0.8562 0.6872
-0.750 -0.0767 0.00600 0.00153 -0.0022 0.8336 0.7081
-0.500 -0.0512 0.00599 0.00151 -0.0014 0.8110 0.7286
-0.250 -0.0256 0.00599 0.00149 -0.0007 0.7892 0.7489
0.000 0.0000 0.00598 0.00149 0.0000 0.7684 0.7684
0.250 0.0256 0.00599 0.00149 0.0007 0.7489 0.7892
0.500 0.0512 0.00598 0.00151 0.0014 0.7286 0.8110
0.750 0.0767 0.00600 0.00153 0.0022 0.7081 0.8336
1.000 0.1024 0.00603 0.00157 0.0029 0.6873 0.8562
1.250 0.1284 0.00607 0.00161 0.0036 0.6652 0.8796
1.500 0.1555 0.00614 0.00167 0.0040 0.6427 0.9022
1.750 0.1845 0.00624 0.00174 0.0040 0.6192 0.9230
2.000 0.2159 0.00636 0.00182 0.0035 0.5951 0.9404
2.250 0.2505 0.00651 0.00191 0.0022 0.5705 0.9532
2.750 0.3221 0.00684 0.00210 -0.0011 0.5177 0.9717
3.000 0.3605 0.00702 0.00221 -0.0034 0.4884 0.9774
3.250 0.3963 0.00724 0.00232 -0.0051 0.4553 0.9843
3.500 0.4349 0.00747 0.00243 -0.0075 0.4139 0.9892
3.750 0.4712 0.00782 0.00258 -0.0096 0.3529 0.9946
4.000 0.5074 0.00856 0.00281 -0.0120 0.2330 0.9993
4.250 0.5288 0.00925 0.00312 -0.0112 0.1514 1.0000
4.500 0.5480 0.00980 0.00344 -0.0098 0.1008 1.0000
4.750 0.5681 0.01029 0.00378 -0.0085 0.0709 1.0000
5.000 0.5890 0.01071 0.00414 -0.0072 0.0553 1.0000
5.250 0.6099 0.01118 0.00458 -0.0059 0.0459 1.0000
5.500 0.6311 0.01164 0.00501 -0.0047 0.0390 1.0000
5.750 0.6523 0.01211 0.00554 -0.0034 0.0349 1.0000
6.000 0.6741 0.01253 0.00598 -0.0023 0.0316 1.0000
6.250 0.6915 0.01354 0.00703 -0.0005 0.0281 1.0000
6.500 0.7142 0.01389 0.00743 0.0005 0.0266 1.0000
6.750 0.7361 0.01433 0.00792 0.0016 0.0245 1.0000
7.000 0.7571 0.01490 0.00851 0.0028 0.0229 1.0000
7.250 0.7718 0.01649 0.01017 0.0049 0.0211 1.0000
7.500 0.7935 0.01711 0.01087 0.0060 0.0204 1.0000
7.750 0.8147 0.01783 0.01168 0.0071 0.0195 1.0000
8.000 0.8360 0.01853 0.01246 0.0082 0.0184 1.0000
8.250 0.8570 0.01923 0.01322 0.0092 0.0174 1.0000
8.500 0.8771 0.02008 0.01411 0.0102 0.0166 1.0000
8.750 0.8925 0.02223 0.01641 0.0118 0.0158 1.0000
9.000 0.9090 0.02417 0.01857 0.0133 0.0154 1.0000
9.250 0.9268 0.02552 0.02013 0.0146 0.0151 1.0000
9.500 0.9424 0.02722 0.02206 0.0161 0.0147 1.0000
9.750 0.9552 0.02927 0.02439 0.0179 0.0143 1.0000
10.000 0.9635 0.03184 0.02729 0.0199 0.0141 1.0000
10.250 0.9652 0.03507 0.03087 0.0223 0.0140 1.0000
10.500 0.9585 0.03891 0.03508 0.0252 0.0140 1.0000
10.750 0.9471 0.04230 0.03877 0.0282 0.0140 1.0000
11.000 0.9302 0.04517 0.04185 0.0315 0.0138 1.0000
11.250 0.9030 0.04961 0.04653 0.0328 0.0140 1.0000
11.500 0.8802 0.05444 0.05155 0.0316 0.0142 1.0000
11.750 0.8560 0.06054 0.05782 0.0279 0.0144 1.0000
12.000 0.8260 0.06911 0.06655 0.0215 0.0145 1.0000
12.250 0.8037 0.07764 0.07518 0.0148 0.0149 1.0000
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Polar data table (+)
Polar graphs
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