RAE 100 AIRFOIL (rae100-il) Xfoil prediction polar at RE=50,000 Ncrit=5
| Details | Polar file |
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Airfoil: RAE 100 AIRFOIL (rae100-il) Reynolds number: 50,000 Max Cl/Cd: 26.08 at α=5.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-rae100-il-50000-n5.txt Download as CSV file: xf-rae100-il-50000-n5.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 = 0.050 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-11.250 -0.7759 0.08501 0.07745 -0.0146 1.0000 0.0610
-11.000 -0.7996 0.07737 0.06975 -0.0199 1.0000 0.0606
-10.750 -0.8237 0.07108 0.06335 -0.0236 1.0000 0.0602
-10.500 -0.8482 0.06597 0.05808 -0.0250 1.0000 0.0601
-10.250 -0.8691 0.06161 0.05350 -0.0245 1.0000 0.0603
-10.000 -0.8835 0.05722 0.04874 -0.0237 1.0000 0.0608
-9.750 -0.8870 0.05335 0.04451 -0.0228 1.0000 0.0616
-9.500 -0.8783 0.05065 0.04171 -0.0219 1.0000 0.0632
-9.250 -0.8693 0.04823 0.03912 -0.0209 1.0000 0.0659
-9.000 -0.8610 0.04542 0.03597 -0.0198 1.0000 0.0690
-8.750 -0.8515 0.04233 0.03232 -0.0185 1.0000 0.0720
-8.500 -0.8354 0.03993 0.02978 -0.0176 1.0000 0.0750
-8.250 -0.8184 0.03814 0.02787 -0.0166 1.0000 0.0797
-8.000 -0.8008 0.03594 0.02515 -0.0155 1.0000 0.0853
-7.750 -0.7811 0.03426 0.02357 -0.0146 1.0000 0.0903
-7.500 -0.7612 0.03260 0.02157 -0.0136 1.0000 0.0982
-7.250 -0.7412 0.03114 0.02019 -0.0127 1.0000 0.1057
-7.000 -0.7207 0.02971 0.01864 -0.0116 1.0000 0.1157
-6.750 -0.7000 0.02840 0.01723 -0.0106 1.0000 0.1277
-6.500 -0.6797 0.02719 0.01602 -0.0096 1.0000 0.1422
-6.250 -0.6597 0.02603 0.01492 -0.0085 1.0000 0.1592
-6.000 -0.6394 0.02496 0.01386 -0.0074 1.0000 0.1812
-5.750 -0.6193 0.02394 0.01295 -0.0063 1.0000 0.2066
-5.500 -0.5991 0.02301 0.01213 -0.0052 1.0000 0.2370
-5.250 -0.5786 0.02219 0.01141 -0.0041 1.0000 0.2723
-5.000 -0.5583 0.02145 0.01081 -0.0029 1.0000 0.3095
-4.750 -0.5376 0.02082 0.01027 -0.0016 1.0000 0.3493
-4.500 -0.5170 0.02026 0.00985 -0.0002 1.0000 0.3885
-4.250 -0.4960 0.01977 0.00946 0.0012 1.0000 0.4286
-4.000 -0.4753 0.01933 0.00912 0.0027 1.0000 0.4669
-3.750 -0.4548 0.01895 0.00882 0.0044 1.0000 0.5050
-3.500 -0.4343 0.01860 0.00858 0.0062 1.0000 0.5410
-3.250 -0.4138 0.01829 0.00835 0.0080 1.0000 0.5770
-3.000 -0.3932 0.01800 0.00812 0.0098 1.0000 0.6126
-2.750 -0.3726 0.01774 0.00796 0.0117 1.0000 0.6467
-2.500 -0.3518 0.01751 0.00782 0.0136 1.0000 0.6808
-2.250 -0.3308 0.01731 0.00769 0.0155 1.0000 0.7149
-2.000 -0.3092 0.01714 0.00760 0.0173 1.0000 0.7494
-1.750 -0.2865 0.01701 0.00753 0.0189 1.0000 0.7849
-1.500 -0.2597 0.01694 0.00753 0.0197 1.0000 0.8199
-1.250 -0.2271 0.01693 0.00754 0.0193 1.0000 0.8557
-1.000 -0.1876 0.01696 0.00757 0.0173 1.0000 0.8908
-0.750 -0.1417 0.01702 0.00759 0.0138 1.0000 0.9236
-0.500 -0.0900 0.01706 0.00759 0.0089 1.0000 0.9527
-0.250 -0.0362 0.01704 0.00756 0.0031 1.0000 0.9798
0.000 0.0000 0.01700 0.00752 0.0000 1.0000 1.0000
0.250 0.0362 0.01704 0.00756 -0.0031 0.9798 1.0000
0.500 0.0900 0.01706 0.00759 -0.0089 0.9527 1.0000
0.750 0.1416 0.01702 0.00759 -0.0138 0.9236 1.0000
1.000 0.1877 0.01696 0.00757 -0.0173 0.8908 1.0000
1.250 0.2271 0.01692 0.00754 -0.0193 0.8557 1.0000
1.500 0.2597 0.01694 0.00753 -0.0197 0.8200 1.0000
1.750 0.2866 0.01701 0.00753 -0.0189 0.7850 1.0000
2.000 0.3093 0.01714 0.00760 -0.0173 0.7496 1.0000
2.250 0.3308 0.01731 0.00769 -0.0155 0.7149 1.0000
2.500 0.3518 0.01751 0.00782 -0.0137 0.6808 1.0000
2.750 0.3726 0.01774 0.00796 -0.0117 0.6467 1.0000
3.000 0.3933 0.01800 0.00812 -0.0098 0.6126 1.0000
3.250 0.4138 0.01829 0.00835 -0.0080 0.5769 1.0000
3.500 0.4344 0.01860 0.00858 -0.0062 0.5410 1.0000
3.750 0.4549 0.01895 0.00882 -0.0044 0.5050 1.0000
4.000 0.4754 0.01933 0.00912 -0.0028 0.4669 1.0000
4.250 0.4961 0.01977 0.00946 -0.0012 0.4285 1.0000
4.500 0.5171 0.02026 0.00985 0.0002 0.3885 1.0000
4.750 0.5377 0.02082 0.01027 0.0016 0.3494 1.0000
5.000 0.5584 0.02145 0.01081 0.0029 0.3095 1.0000
5.250 0.5787 0.02219 0.01141 0.0041 0.2723 1.0000
5.500 0.5991 0.02301 0.01213 0.0052 0.2370 1.0000
5.750 0.6193 0.02394 0.01295 0.0063 0.2066 1.0000
6.000 0.6394 0.02495 0.01386 0.0074 0.1812 1.0000
6.250 0.6596 0.02603 0.01492 0.0085 0.1592 1.0000
6.500 0.6797 0.02719 0.01603 0.0096 0.1423 1.0000
6.750 0.6999 0.02840 0.01723 0.0106 0.1277 1.0000
7.000 0.7206 0.02971 0.01864 0.0117 0.1158 1.0000
7.250 0.7411 0.03114 0.02019 0.0127 0.1057 1.0000
7.500 0.7610 0.03260 0.02156 0.0136 0.0983 1.0000
7.750 0.7810 0.03426 0.02356 0.0147 0.0903 1.0000
8.000 0.8007 0.03594 0.02515 0.0155 0.0852 1.0000
8.250 0.8182 0.03812 0.02785 0.0167 0.0797 1.0000
8.500 0.8353 0.03991 0.02977 0.0176 0.0750 1.0000
8.750 0.8514 0.04230 0.03229 0.0185 0.0720 1.0000
9.000 0.8609 0.04540 0.03595 0.0198 0.0690 1.0000
9.250 0.8691 0.04820 0.03910 0.0209 0.0659 1.0000
9.500 0.8781 0.05064 0.04171 0.0219 0.0633 1.0000
9.750 0.8869 0.05332 0.04449 0.0228 0.0616 1.0000
10.000 0.8834 0.05720 0.04871 0.0238 0.0608 1.0000
10.250 0.8689 0.06160 0.05349 0.0245 0.0603 1.0000
10.500 0.8494 0.06587 0.05797 0.0251 0.0602 1.0000
10.750 0.8243 0.07100 0.06327 0.0237 0.0603 1.0000
11.000 0.7997 0.07737 0.06974 0.0199 0.0606 1.0000
11.250 0.7767 0.08490 0.07733 0.0147 0.0610 1.0000
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Polar data table (+)
Polar graphs
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