RAE 101 AIRFOIL (rae101-il) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: RAE 101 AIRFOIL (rae101-il) Reynolds number: 200,000 Max Cl/Cd: 49.65 at α=4.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-rae101-il-200000.txt Download as CSV file: xf-rae101-il-200000.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 = 0.200 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.500 -0.8410 0.05365 0.04905 -0.0242 1.0000 0.0364
-9.250 -0.8464 0.04864 0.04383 -0.0228 1.0000 0.0354
-9.000 -0.8474 0.04473 0.03957 -0.0209 1.0000 0.0354
-8.750 -0.8444 0.04081 0.03525 -0.0189 1.0000 0.0351
-8.500 -0.8381 0.03685 0.03088 -0.0169 1.0000 0.0347
-8.250 -0.8274 0.03330 0.02691 -0.0149 1.0000 0.0343
-8.000 -0.8128 0.03034 0.02355 -0.0131 1.0000 0.0345
-7.750 -0.7950 0.02796 0.02080 -0.0115 1.0000 0.0353
-7.500 -0.7748 0.02656 0.01908 -0.0102 1.0000 0.0367
-7.250 -0.7555 0.02376 0.01607 -0.0090 1.0000 0.0388
-7.000 -0.7341 0.02221 0.01448 -0.0080 1.0000 0.0409
-6.750 -0.7121 0.02094 0.01309 -0.0069 1.0000 0.0433
-6.500 -0.6900 0.01989 0.01192 -0.0058 1.0000 0.0464
-6.250 -0.6701 0.01842 0.01043 -0.0044 1.0000 0.0504
-6.000 -0.6498 0.01748 0.00950 -0.0031 1.0000 0.0551
-5.750 -0.6293 0.01663 0.00857 -0.0016 1.0000 0.0604
-5.500 -0.6109 0.01564 0.00763 0.0000 1.0000 0.0694
-5.250 -0.5924 0.01470 0.00674 0.0016 1.0000 0.0805
-5.000 -0.5737 0.01383 0.00595 0.0032 1.0000 0.1014
-4.750 -0.5572 0.01269 0.00516 0.0050 1.0000 0.1546
-4.500 -0.5448 0.01124 0.00455 0.0071 1.0000 0.3009
-4.250 -0.5275 0.01062 0.00434 0.0089 1.0000 0.4089
-4.000 -0.5077 0.01032 0.00420 0.0105 1.0000 0.4644
-3.750 -0.4871 0.01011 0.00409 0.0119 1.0000 0.5047
-3.500 -0.4665 0.00995 0.00400 0.0134 1.0000 0.5377
-3.250 -0.4460 0.00982 0.00392 0.0149 1.0000 0.5672
-3.000 -0.4258 0.00972 0.00387 0.0163 1.0000 0.5940
-2.750 -0.4060 0.00964 0.00385 0.0179 1.0000 0.6200
-2.500 -0.3867 0.00960 0.00387 0.0194 1.0000 0.6456
-2.250 -0.3637 0.00957 0.00392 0.0202 0.9985 0.6717
-2.000 -0.3200 0.00953 0.00396 0.0169 0.9890 0.7030
-1.750 -0.2773 0.00945 0.00398 0.0139 0.9796 0.7324
-1.500 -0.2339 0.00936 0.00396 0.0109 0.9708 0.7590
-1.250 -0.1934 0.00923 0.00391 0.0085 0.9600 0.7836
-1.000 -0.1514 0.00911 0.00385 0.0059 0.9502 0.8078
-0.750 -0.1092 0.00898 0.00380 0.0034 0.9402 0.8290
-0.500 -0.0712 0.00888 0.00374 0.0018 0.9266 0.8516
-0.250 -0.0347 0.00882 0.00372 0.0007 0.9105 0.8717
0.000 0.0000 0.00879 0.00371 0.0000 0.8914 0.8914
0.250 0.0347 0.00882 0.00372 -0.0007 0.8717 0.9105
0.500 0.0711 0.00888 0.00374 -0.0018 0.8515 0.9265
0.750 0.1092 0.00898 0.00380 -0.0034 0.8291 0.9402
1.000 0.1513 0.00911 0.00385 -0.0059 0.8078 0.9502
1.250 0.1934 0.00923 0.00391 -0.0085 0.7835 0.9600
1.500 0.2339 0.00935 0.00396 -0.0109 0.7592 0.9708
1.750 0.2774 0.00945 0.00398 -0.0139 0.7325 0.9796
2.000 0.3200 0.00952 0.00396 -0.0169 0.7032 0.9891
2.250 0.3636 0.00956 0.00392 -0.0202 0.6717 0.9986
2.500 0.3865 0.00960 0.00387 -0.0194 0.6457 1.0000
2.750 0.4058 0.00964 0.00385 -0.0178 0.6199 1.0000
3.000 0.4256 0.00971 0.00387 -0.0163 0.5941 1.0000
3.250 0.4458 0.00982 0.00392 -0.0148 0.5674 1.0000
3.500 0.4663 0.00995 0.00400 -0.0134 0.5380 1.0000
3.750 0.4869 0.01011 0.00409 -0.0119 0.5045 1.0000
4.000 0.5075 0.01031 0.00420 -0.0104 0.4649 1.0000
4.250 0.5273 0.01062 0.00434 -0.0089 0.4090 1.0000
4.500 0.5446 0.01123 0.00455 -0.0071 0.3012 1.0000
4.750 0.5571 0.01269 0.00516 -0.0050 0.1543 1.0000
5.000 0.5736 0.01383 0.00595 -0.0032 0.1015 1.0000
5.250 0.5923 0.01471 0.00674 -0.0016 0.0807 1.0000
5.500 0.6108 0.01564 0.00763 0.0001 0.0693 1.0000
5.750 0.6292 0.01663 0.00856 0.0016 0.0604 1.0000
6.000 0.6498 0.01748 0.00950 0.0031 0.0551 1.0000
6.250 0.6701 0.01843 0.01043 0.0044 0.0504 1.0000
6.500 0.6900 0.01989 0.01192 0.0058 0.0464 1.0000
6.750 0.7121 0.02094 0.01309 0.0069 0.0433 1.0000
7.000 0.7341 0.02223 0.01449 0.0080 0.0410 1.0000
7.250 0.7556 0.02375 0.01606 0.0090 0.0389 1.0000
7.500 0.7748 0.02665 0.01917 0.0101 0.0367 1.0000
7.750 0.7951 0.02796 0.02080 0.0115 0.0353 1.0000
8.000 0.8129 0.03035 0.02355 0.0131 0.0346 1.0000
8.250 0.8276 0.03328 0.02688 0.0149 0.0344 1.0000
8.500 0.8383 0.03685 0.03088 0.0168 0.0347 1.0000
8.750 0.8447 0.04080 0.03524 0.0188 0.0352 1.0000
9.000 0.8477 0.04475 0.03958 0.0208 0.0355 1.0000
9.250 0.8467 0.04895 0.04410 0.0226 0.0358 1.0000
9.500 0.8416 0.05332 0.04874 0.0242 0.0362 1.0000
12.250 0.5664 0.12453 0.12116 -0.0012 0.0597 1.0000
12.500 0.5552 0.12989 0.12649 -0.0053 0.0594 1.0000
|
Polar data table (+)
Polar graphs
<< Back to RAE 101 AIRFOIL (rae101-il)