RAE 100 AIRFOIL (rae100-il) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
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Airfoil: RAE 100 AIRFOIL (rae100-il) Reynolds number: 200,000 Max Cl/Cd: 46.11 at α=4.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-rae100-il-200000.txt Download as CSV file: xf-rae100-il-200000.csv |
XFOIL Version 6.96
Calculated polar for: RAE 100 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
------ -------- --------- --------- -------- -------- --------
-12.500 -0.7609 0.10135 0.09757 -0.0020 1.0000 0.0414
-12.250 -0.7712 0.09463 0.09088 -0.0044 1.0000 0.0407
-12.000 -0.8118 0.08070 0.07690 -0.0140 1.0000 0.0393
-11.750 -0.8592 0.06815 0.06418 -0.0244 1.0000 0.0382
-11.500 -0.8993 0.05989 0.05569 -0.0301 1.0000 0.0372
-11.250 -0.9402 0.05374 0.04923 -0.0312 1.0000 0.0364
-11.000 -0.9708 0.04944 0.04461 -0.0285 1.0000 0.0360
-10.750 -0.9868 0.04528 0.04005 -0.0262 1.0000 0.0358
-10.500 -0.9864 0.04229 0.03677 -0.0244 1.0000 0.0362
-10.250 -0.9819 0.03948 0.03369 -0.0227 1.0000 0.0368
-10.000 -0.9739 0.03691 0.03082 -0.0211 1.0000 0.0376
-9.750 -0.9626 0.03473 0.02835 -0.0195 1.0000 0.0388
-9.500 -0.9495 0.03279 0.02605 -0.0180 1.0000 0.0404
-9.250 -0.9342 0.03124 0.02411 -0.0164 1.0000 0.0414
-9.000 -0.9182 0.02799 0.02069 -0.0153 1.0000 0.0430
-8.750 -0.8981 0.02652 0.01916 -0.0144 1.0000 0.0447
-8.500 -0.8775 0.02539 0.01791 -0.0134 1.0000 0.0472
-8.250 -0.8563 0.02437 0.01667 -0.0123 1.0000 0.0499
-8.000 -0.8360 0.02250 0.01471 -0.0113 1.0000 0.0526
-7.750 -0.8145 0.02143 0.01364 -0.0104 1.0000 0.0557
-7.500 -0.7922 0.02062 0.01271 -0.0094 1.0000 0.0598
-7.250 -0.7717 0.01932 0.01141 -0.0083 1.0000 0.0644
-7.000 -0.7498 0.01853 0.01061 -0.0073 1.0000 0.0696
-6.750 -0.7290 0.01754 0.00959 -0.0061 1.0000 0.0761
-6.500 -0.7068 0.01689 0.00891 -0.0050 1.0000 0.0841
-6.250 -0.6863 0.01601 0.00810 -0.0038 1.0000 0.0939
-6.000 -0.6653 0.01526 0.00740 -0.0026 1.0000 0.1066
-5.750 -0.6440 0.01457 0.00676 -0.0014 1.0000 0.1234
-5.500 -0.6233 0.01386 0.00618 -0.0002 1.0000 0.1453
-5.250 -0.6022 0.01322 0.00570 0.0009 1.0000 0.1736
-5.000 -0.5811 0.01264 0.00529 0.0021 1.0000 0.2077
-4.750 -0.5596 0.01214 0.00493 0.0032 1.0000 0.2454
-4.500 -0.5382 0.01168 0.00465 0.0043 1.0000 0.2842
-4.250 -0.5166 0.01129 0.00441 0.0055 1.0000 0.3232
-4.000 -0.4950 0.01094 0.00421 0.0067 1.0000 0.3616
-3.750 -0.4735 0.01063 0.00403 0.0079 1.0000 0.3990
-3.500 -0.4521 0.01036 0.00390 0.0092 1.0000 0.4354
-3.250 -0.4308 0.01012 0.00379 0.0105 1.0000 0.4711
-3.000 -0.4098 0.00992 0.00370 0.0118 1.0000 0.5058
-2.750 -0.3892 0.00973 0.00365 0.0133 1.0000 0.5393
-2.500 -0.3690 0.00958 0.00363 0.0148 1.0000 0.5724
-2.250 -0.3493 0.00947 0.00365 0.0163 1.0000 0.6054
-2.000 -0.3303 0.00939 0.00370 0.0180 1.0000 0.6379
-1.750 -0.2985 0.00934 0.00379 0.0170 0.9952 0.6746
-1.500 -0.2536 0.00923 0.00385 0.0137 0.9850 0.7126
-1.250 -0.2101 0.00911 0.00386 0.0107 0.9738 0.7496
-0.750 -0.1229 0.00884 0.00385 0.0054 0.9520 0.8184
-0.500 -0.0808 0.00873 0.00383 0.0033 0.9388 0.8494
-0.250 -0.0412 0.00866 0.00383 0.0018 0.9214 0.8776
0.000 0.0000 0.00864 0.00384 0.0000 0.9011 0.9010
0.250 0.0412 0.00866 0.00383 -0.0018 0.8777 0.9213
0.500 0.0808 0.00873 0.00384 -0.0033 0.8494 0.9389
0.750 0.1229 0.00884 0.00385 -0.0054 0.8183 0.9519
1.000 0.1672 0.00898 0.00386 -0.0081 0.7852 0.9625
1.250 0.2101 0.00911 0.00386 -0.0108 0.7496 0.9738
1.500 0.2536 0.00924 0.00385 -0.0137 0.7126 0.9849
1.750 0.2984 0.00934 0.00379 -0.0170 0.6745 0.9952
2.000 0.3304 0.00939 0.00370 -0.0180 0.6376 1.0000
2.250 0.3495 0.00947 0.00365 -0.0164 0.6052 1.0000
2.500 0.3692 0.00958 0.00363 -0.0148 0.5725 1.0000
2.750 0.3894 0.00973 0.00365 -0.0133 0.5394 1.0000
3.000 0.4100 0.00992 0.00370 -0.0119 0.5059 1.0000
3.250 0.4310 0.01012 0.00379 -0.0105 0.4712 1.0000
3.500 0.4523 0.01036 0.00390 -0.0092 0.4352 1.0000
3.750 0.4736 0.01063 0.00403 -0.0079 0.3989 1.0000
4.000 0.4952 0.01094 0.00421 -0.0067 0.3615 1.0000
4.250 0.5168 0.01129 0.00441 -0.0055 0.3231 1.0000
4.500 0.5384 0.01168 0.00465 -0.0044 0.2841 1.0000
4.750 0.5598 0.01214 0.00493 -0.0032 0.2454 1.0000
5.000 0.5813 0.01264 0.00529 -0.0021 0.2075 1.0000
5.250 0.6024 0.01322 0.00570 -0.0010 0.1736 1.0000
5.500 0.6234 0.01386 0.00619 0.0002 0.1452 1.0000
5.750 0.6442 0.01457 0.00676 0.0014 0.1234 1.0000
6.000 0.6654 0.01526 0.00740 0.0026 0.1065 1.0000
6.250 0.6864 0.01600 0.00810 0.0038 0.0938 1.0000
6.500 0.7068 0.01689 0.00891 0.0050 0.0841 1.0000
6.750 0.7290 0.01753 0.00959 0.0061 0.0760 1.0000
7.000 0.7497 0.01853 0.01061 0.0073 0.0696 1.0000
7.250 0.7716 0.01933 0.01142 0.0083 0.0644 1.0000
7.500 0.7921 0.02065 0.01274 0.0094 0.0599 1.0000
7.750 0.8144 0.02145 0.01365 0.0104 0.0558 1.0000
8.000 0.8358 0.02250 0.01471 0.0114 0.0526 1.0000
8.250 0.8562 0.02436 0.01666 0.0124 0.0499 1.0000
8.500 0.8774 0.02538 0.01790 0.0135 0.0472 1.0000
8.750 0.8979 0.02655 0.01919 0.0145 0.0448 1.0000
9.000 0.9180 0.02798 0.02068 0.0154 0.0430 1.0000
9.250 0.9340 0.03127 0.02414 0.0164 0.0415 1.0000
9.500 0.9494 0.03272 0.02598 0.0180 0.0403 1.0000
9.750 0.9623 0.03477 0.02839 0.0196 0.0390 1.0000
10.000 0.9734 0.03696 0.03088 0.0212 0.0377 1.0000
10.250 0.9820 0.03941 0.03361 0.0227 0.0367 1.0000
10.500 0.9851 0.04241 0.03692 0.0245 0.0364 1.0000
10.750 0.9868 0.04522 0.03999 0.0262 0.0358 1.0000
11.000 0.9761 0.04891 0.04400 0.0283 0.0358 1.0000
11.250 0.9404 0.05370 0.04918 0.0313 0.0364 1.0000
11.500 0.8979 0.06009 0.05590 0.0300 0.0375 1.0000
11.750 0.8590 0.06816 0.06419 0.0244 0.0381 1.0000
12.000 0.8112 0.08086 0.07707 0.0138 0.0393 1.0000
12.250 0.7723 0.09445 0.09070 0.0045 0.0407 1.0000
12.500 0.7612 0.10139 0.09761 0.0018 0.0414 1.0000
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Polar data table (+)
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