RAE 104 AIRFOIL (rae104-il) Xfoil prediction polar at RE=200,000 Ncrit=5
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Airfoil: RAE 104 AIRFOIL (rae104-il) Reynolds number: 200,000 Max Cl/Cd: 36.66 at α=3.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-rae104-il-200000-n5.txt Download as CSV file: xf-rae104-il-200000-n5.csv |
XFOIL Version 6.96
Calculated polar for: RAE 104 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 = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-11.250 -0.6610 0.08810 0.08449 -0.0245 1.0000 0.0137
-11.000 -0.6859 0.07511 0.07149 -0.0339 1.0000 0.0133
-10.750 -0.7166 0.06623 0.06250 -0.0392 1.0000 0.0129
-10.500 -0.7423 0.06042 0.05655 -0.0408 1.0000 0.0128
-10.250 -0.7679 0.05555 0.05149 -0.0399 1.0000 0.0126
-10.000 -0.7895 0.05176 0.04750 -0.0371 1.0000 0.0126
-9.750 -0.8077 0.04840 0.04392 -0.0330 1.0000 0.0126
-9.500 -0.8200 0.04467 0.03988 -0.0294 1.0000 0.0127
-9.250 -0.8276 0.04097 0.03582 -0.0258 1.0000 0.0127
-9.000 -0.8296 0.03761 0.03209 -0.0224 1.0000 0.0129
-8.750 -0.8267 0.03462 0.02871 -0.0193 1.0000 0.0131
-8.500 -0.8198 0.03196 0.02564 -0.0165 1.0000 0.0133
-8.250 -0.8095 0.02963 0.02296 -0.0139 1.0000 0.0138
-8.000 -0.7965 0.02756 0.02056 -0.0116 1.0000 0.0143
-7.750 -0.7812 0.02580 0.01851 -0.0096 1.0000 0.0148
-7.500 -0.7645 0.02442 0.01688 -0.0077 1.0000 0.0155
-7.250 -0.7491 0.02292 0.01525 -0.0058 1.0000 0.0167
-7.000 -0.7333 0.02209 0.01437 -0.0040 1.0000 0.0180
-6.750 -0.7174 0.02114 0.01334 -0.0020 1.0000 0.0189
-6.500 -0.7026 0.02019 0.01226 0.0003 1.0000 0.0200
-6.250 -0.6817 0.01921 0.01116 0.0013 0.9982 0.0213
-6.000 -0.6509 0.01829 0.01012 0.0002 0.9929 0.0230
-5.750 -0.6223 0.01726 0.00908 -0.0007 0.9869 0.0264
-5.500 -0.5915 0.01652 0.00825 -0.0018 0.9816 0.0299
-5.250 -0.5625 0.01575 0.00737 -0.0025 0.9748 0.0335
-5.000 -0.5303 0.01512 0.00672 -0.0039 0.9703 0.0416
-4.750 -0.5019 0.01449 0.00609 -0.0044 0.9630 0.0513
-4.500 -0.4710 0.01390 0.00554 -0.0055 0.9576 0.0708
-4.250 -0.4432 0.01327 0.00505 -0.0060 0.9509 0.1071
-4.000 -0.4165 0.01247 0.00460 -0.0065 0.9443 0.1824
-3.750 -0.3935 0.01154 0.00417 -0.0063 0.9370 0.2990
-3.500 -0.3745 0.01045 0.00378 -0.0052 0.9293 0.4572
-3.250 -0.3573 0.00975 0.00379 -0.0029 0.9210 0.6144
-3.000 -0.3302 0.00964 0.00379 -0.0025 0.9152 0.6825
-2.750 -0.3032 0.00962 0.00377 -0.0021 0.9092 0.7174
-2.500 -0.2763 0.00961 0.00376 -0.0016 0.9032 0.7432
-2.250 -0.2469 0.00962 0.00374 -0.0017 0.8989 0.7644
-2.000 -0.2216 0.00965 0.00375 -0.0009 0.8922 0.7822
-1.750 -0.1937 0.00967 0.00374 -0.0007 0.8870 0.7968
-1.500 -0.1657 0.00968 0.00373 -0.0005 0.8819 0.8083
-1.250 -0.1388 0.00969 0.00371 -0.0002 0.8762 0.8178
-1.000 -0.1109 0.00968 0.00365 -0.0002 0.8715 0.8265
-0.750 -0.0830 0.00969 0.00365 -0.0001 0.8660 0.8327
-0.500 -0.0560 0.00968 0.00363 0.0001 0.8606 0.8395
-0.250 -0.0269 0.00967 0.00359 -0.0002 0.8563 0.8447
0.000 0.0000 0.00968 0.00362 0.0000 0.8502 0.8502
0.250 0.0269 0.00967 0.00359 0.0002 0.8447 0.8563
0.500 0.0560 0.00968 0.00363 -0.0001 0.8395 0.8606
0.750 0.0830 0.00969 0.00365 0.0001 0.8326 0.8660
1.000 0.1109 0.00968 0.00365 0.0002 0.8265 0.8715
1.250 0.1388 0.00969 0.00371 0.0002 0.8178 0.8762
1.500 0.1657 0.00968 0.00373 0.0005 0.8082 0.8819
1.750 0.1937 0.00967 0.00373 0.0007 0.7969 0.8869
2.000 0.2216 0.00965 0.00375 0.0009 0.7823 0.8922
2.250 0.2469 0.00962 0.00374 0.0017 0.7644 0.8989
2.500 0.2763 0.00961 0.00376 0.0016 0.7431 0.9032
2.750 0.3032 0.00962 0.00377 0.0021 0.7172 0.9092
3.000 0.3302 0.00964 0.00378 0.0025 0.6824 0.9152
3.250 0.3574 0.00975 0.00380 0.0029 0.6154 0.9210
3.500 0.3745 0.01045 0.00378 0.0052 0.4576 0.9293
3.750 0.3936 0.01153 0.00417 0.0063 0.3002 0.9370
4.000 0.4166 0.01247 0.00460 0.0065 0.1834 0.9443
4.250 0.4431 0.01328 0.00505 0.0060 0.1071 0.9509
4.500 0.4710 0.01390 0.00554 0.0055 0.0709 0.9576
4.750 0.5019 0.01449 0.00609 0.0044 0.0514 0.9630
5.000 0.5302 0.01512 0.00672 0.0039 0.0418 0.9703
5.250 0.5627 0.01573 0.00736 0.0025 0.0339 0.9748
5.500 0.5915 0.01651 0.00825 0.0018 0.0300 0.9815
5.750 0.6223 0.01725 0.00908 0.0007 0.0265 0.9868
6.000 0.6510 0.01827 0.01010 -0.0002 0.0230 0.9929
6.250 0.6817 0.01923 0.01118 -0.0012 0.0213 0.9982
6.500 0.7026 0.02019 0.01226 -0.0003 0.0200 1.0000
6.750 0.7175 0.02114 0.01333 0.0020 0.0189 1.0000
7.000 0.7333 0.02207 0.01436 0.0040 0.0179 1.0000
7.250 0.7490 0.02289 0.01520 0.0058 0.0166 1.0000
7.500 0.7645 0.02442 0.01688 0.0077 0.0156 1.0000
7.750 0.7813 0.02577 0.01848 0.0096 0.0148 1.0000
8.000 0.7966 0.02755 0.02054 0.0116 0.0143 1.0000
8.250 0.8096 0.02962 0.02294 0.0139 0.0137 1.0000
8.500 0.8199 0.03195 0.02563 0.0165 0.0134 1.0000
8.750 0.8268 0.03459 0.02867 0.0193 0.0130 1.0000
9.000 0.8297 0.03757 0.03204 0.0224 0.0128 1.0000
9.250 0.8279 0.04092 0.03577 0.0258 0.0127 1.0000
9.500 0.8207 0.04458 0.03978 0.0294 0.0127 1.0000
9.750 0.8072 0.04852 0.04404 0.0331 0.0127 1.0000
10.000 0.7886 0.05192 0.04766 0.0371 0.0127 1.0000
10.250 0.7652 0.05600 0.05195 0.0400 0.0128 1.0000
10.500 0.7419 0.06054 0.05667 0.0407 0.0129 1.0000
10.750 0.7150 0.06660 0.06289 0.0390 0.0130 1.0000
11.000 0.6876 0.07481 0.07118 0.0341 0.0133 1.0000
11.250 0.6625 0.08783 0.08422 0.0245 0.0138 1.0000
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Polar data table (+)
Polar graphs
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