RAE 104 AIRFOIL (rae104-il) Xfoil prediction polar at RE=100,000 Ncrit=5
| Details | Polar file |
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Airfoil: RAE 104 AIRFOIL (rae104-il) Reynolds number: 100,000 Max Cl/Cd: 33.38 at α=3.5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-rae104-il-100000-n5.txt Download as CSV file: xf-rae104-il-100000-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 = 0.100 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-11.250 -0.6448 0.09247 0.08733 -0.0247 1.0000 0.0225
-11.000 -0.6556 0.08474 0.07961 -0.0297 1.0000 0.0222
-10.750 -0.6727 0.07754 0.07236 -0.0346 1.0000 0.0220
-10.500 -0.6914 0.07176 0.06651 -0.0377 1.0000 0.0218
-10.250 -0.7104 0.06701 0.06166 -0.0390 1.0000 0.0217
-10.000 -0.7293 0.06300 0.05752 -0.0387 1.0000 0.0216
-9.750 -0.7468 0.05963 0.05400 -0.0367 1.0000 0.0215
-9.500 -0.7635 0.05637 0.05055 -0.0336 1.0000 0.0215
-9.250 -0.7746 0.05289 0.04681 -0.0307 1.0000 0.0215
-9.000 -0.7821 0.04938 0.04297 -0.0277 1.0000 0.0217
-8.750 -0.7852 0.04611 0.03930 -0.0246 1.0000 0.0221
-8.500 -0.7844 0.04330 0.03600 -0.0214 1.0000 0.0227
-8.250 -0.7813 0.04015 0.03262 -0.0190 1.0000 0.0238
-8.000 -0.7722 0.03806 0.03033 -0.0168 1.0000 0.0248
-7.750 -0.7617 0.03570 0.02767 -0.0145 1.0000 0.0254
-7.500 -0.7490 0.03339 0.02502 -0.0122 1.0000 0.0261
-7.250 -0.7340 0.03125 0.02258 -0.0102 1.0000 0.0270
-7.000 -0.7169 0.02927 0.02030 -0.0084 1.0000 0.0283
-6.750 -0.6986 0.02769 0.01842 -0.0066 1.0000 0.0302
-6.500 -0.6801 0.02623 0.01672 -0.0050 1.0000 0.0327
-6.250 -0.6631 0.02471 0.01519 -0.0033 1.0000 0.0349
-6.000 -0.6465 0.02354 0.01396 -0.0014 1.0000 0.0371
-5.750 -0.6310 0.02251 0.01284 0.0008 1.0000 0.0400
-5.500 -0.6171 0.02163 0.01187 0.0032 1.0000 0.0444
-5.250 -0.6052 0.02078 0.01103 0.0058 1.0000 0.0488
-5.000 -0.5930 0.02004 0.01016 0.0085 1.0000 0.0538
-4.750 -0.5814 0.01933 0.00946 0.0111 1.0000 0.0617
-4.500 -0.5594 0.01854 0.00868 0.0117 0.9968 0.0757
-4.250 -0.5307 0.01763 0.00790 0.0109 0.9907 0.1077
-4.000 -0.5044 0.01648 0.00726 0.0102 0.9847 0.1960
-3.750 -0.4861 0.01496 0.00670 0.0109 0.9776 0.3845
-3.500 -0.4677 0.01402 0.00697 0.0131 0.9718 0.6265
-3.250 -0.4417 0.01409 0.00712 0.0141 0.9658 0.7136
-3.000 -0.4134 0.01427 0.00728 0.0147 0.9603 0.7595
-2.750 -0.3813 0.01452 0.00743 0.0146 0.9565 0.7942
-2.500 -0.3529 0.01478 0.00762 0.0154 0.9514 0.8207
-2.250 -0.3226 0.01501 0.00775 0.0155 0.9469 0.8428
-2.000 -0.2858 0.01524 0.00785 0.0142 0.9440 0.8583
-1.750 -0.2495 0.01535 0.00785 0.0126 0.9409 0.8703
-1.500 -0.2140 0.01547 0.00788 0.0111 0.9365 0.8775
-1.250 -0.1810 0.01547 0.00780 0.0098 0.9320 0.8857
-1.000 -0.1402 0.01552 0.00776 0.0071 0.9293 0.8905
-0.750 -0.1086 0.01551 0.00769 0.0060 0.9244 0.8976
-0.500 -0.0726 0.01554 0.00769 0.0041 0.9198 0.9021
-0.250 -0.0342 0.01554 0.00766 0.0016 0.9162 0.9064
0.000 0.0000 0.01551 0.00761 0.0000 0.9122 0.9121
0.250 0.0342 0.01554 0.00766 -0.0016 0.9063 0.9162
0.500 0.0726 0.01554 0.00769 -0.0041 0.9021 0.9198
0.750 0.1085 0.01551 0.00769 -0.0060 0.8976 0.9244
1.000 0.1402 0.01552 0.00776 -0.0071 0.8905 0.9293
1.250 0.1809 0.01548 0.00780 -0.0098 0.8858 0.9320
1.500 0.2140 0.01547 0.00788 -0.0111 0.8776 0.9365
1.750 0.2495 0.01535 0.00785 -0.0126 0.8703 0.9409
2.000 0.2857 0.01524 0.00788 -0.0142 0.8583 0.9440
2.250 0.3226 0.01501 0.00775 -0.0155 0.8428 0.9469
2.500 0.3529 0.01478 0.00762 -0.0153 0.8207 0.9514
2.750 0.3812 0.01452 0.00743 -0.0147 0.7942 0.9565
3.000 0.4134 0.01427 0.00728 -0.0147 0.7596 0.9603
3.250 0.4416 0.01409 0.00712 -0.0141 0.7128 0.9659
3.500 0.4677 0.01401 0.00696 -0.0131 0.6249 0.9718
3.750 0.4862 0.01495 0.00670 -0.0109 0.3853 0.9776
4.000 0.5045 0.01647 0.00726 -0.0102 0.1961 0.9847
4.250 0.5307 0.01763 0.00790 -0.0109 0.1077 0.9907
4.500 0.5594 0.01854 0.00868 -0.0117 0.0757 0.9968
4.750 0.5814 0.01932 0.00946 -0.0111 0.0618 1.0000
5.000 0.5930 0.02004 0.01016 -0.0085 0.0540 1.0000
5.250 0.6052 0.02077 0.01103 -0.0058 0.0491 1.0000
5.500 0.6172 0.02162 0.01186 -0.0032 0.0445 1.0000
5.750 0.6310 0.02252 0.01284 -0.0008 0.0403 1.0000
6.000 0.6465 0.02353 0.01396 0.0014 0.0372 1.0000
6.250 0.6631 0.02471 0.01519 0.0033 0.0349 1.0000
6.500 0.6801 0.02624 0.01673 0.0050 0.0327 1.0000
6.750 0.6986 0.02769 0.01841 0.0066 0.0303 1.0000
7.000 0.7169 0.02927 0.02031 0.0084 0.0282 1.0000
7.250 0.7340 0.03127 0.02259 0.0102 0.0271 1.0000
7.500 0.7491 0.03338 0.02501 0.0122 0.0261 1.0000
7.750 0.7618 0.03569 0.02765 0.0145 0.0254 1.0000
8.000 0.7723 0.03804 0.03032 0.0168 0.0248 1.0000
8.250 0.7813 0.04009 0.03253 0.0189 0.0236 1.0000
8.500 0.7846 0.04340 0.03607 0.0213 0.0228 1.0000
8.750 0.7852 0.04610 0.03929 0.0246 0.0221 1.0000
9.000 0.7817 0.04941 0.04301 0.0277 0.0216 1.0000
9.250 0.7746 0.05289 0.04681 0.0307 0.0215 1.0000
9.500 0.7643 0.05629 0.05046 0.0335 0.0215 1.0000
9.750 0.7477 0.05957 0.05393 0.0366 0.0215 1.0000
10.000 0.7301 0.06295 0.05746 0.0386 0.0216 1.0000
10.250 0.7111 0.06697 0.06161 0.0390 0.0217 1.0000
10.500 0.6921 0.07170 0.06645 0.0377 0.0218 1.0000
10.750 0.6730 0.07756 0.07239 0.0345 0.0220 1.0000
11.000 0.6569 0.08454 0.07941 0.0298 0.0222 1.0000
11.250 0.6460 0.09223 0.08709 0.0248 0.0225 1.0000
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Polar data table (+)
Polar graphs
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