DU 86-137/25 AIRFOIL (du861372-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
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Airfoil: DU 86-137/25 AIRFOIL (du861372-il) Reynolds number: 100,000 Max Cl/Cd: 33.41 at α=4.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-du861372-il-100000.txt Download as CSV file: xf-du861372-il-100000.csv |
XFOIL Version 6.96
Calculated polar for: DU 86-137/25 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 = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.750 -0.8082 0.04972 0.04200 -0.0257 0.4160 0.1091
-8.500 -0.8107 0.04399 0.03498 -0.0210 0.4161 0.0903
-8.250 -0.7837 0.04011 0.03093 -0.0209 0.4162 0.0846
-8.000 -0.7651 0.03793 0.02788 -0.0184 0.4162 0.0776
-7.750 -0.7340 0.03496 0.02476 -0.0188 0.4162 0.0749
-7.500 -0.7026 0.03254 0.02210 -0.0190 0.4164 0.0721
-7.250 -0.6694 0.03065 0.01997 -0.0193 0.4165 0.0698
-7.000 -0.6361 0.02909 0.01830 -0.0198 0.4170 0.0696
-6.750 -0.6037 0.02773 0.01689 -0.0202 0.4176 0.0698
-6.500 -0.5737 0.02656 0.01570 -0.0203 0.4184 0.0693
-6.250 -0.5493 0.02555 0.01468 -0.0196 0.4193 0.0690
-6.000 -0.5300 0.02465 0.01376 -0.0182 0.4201 0.0690
-5.750 -0.5130 0.02382 0.01287 -0.0166 0.4212 0.0694
-5.500 -0.4960 0.02306 0.01204 -0.0150 0.4223 0.0703
-5.250 -0.4782 0.02235 0.01124 -0.0135 0.4235 0.0718
-5.000 -0.4592 0.02171 0.01050 -0.0122 0.4248 0.0744
-4.750 -0.4402 0.02103 0.00979 -0.0108 0.4262 0.0803
-4.500 -0.4249 0.02003 0.00918 -0.0091 0.4276 0.1390
-4.250 -0.4069 0.01927 0.00861 -0.0077 0.4291 0.1916
-4.000 -0.4029 0.01762 0.00796 -0.0046 0.4306 0.3726
-3.750 -0.3901 0.01679 0.00777 -0.0020 0.4322 0.4954
-3.500 -0.3674 0.01671 0.00808 -0.0003 0.4343 0.5771
-3.250 -0.3416 0.01676 0.00816 0.0005 0.4369 0.6116
-3.000 -0.3161 0.01675 0.00812 0.0012 0.4397 0.6342
-2.750 -0.2909 0.01674 0.00806 0.0018 0.4427 0.6522
-2.500 -0.2658 0.01675 0.00802 0.0024 0.4458 0.6691
-2.250 -0.2406 0.01689 0.00815 0.0032 0.4491 0.6892
-2.000 -0.2157 0.01687 0.00809 0.0037 0.4530 0.7034
-1.750 -0.1894 0.01686 0.00805 0.0041 0.4574 0.7115
-1.500 -0.1641 0.01692 0.00807 0.0047 0.4620 0.7235
-1.250 -0.1393 0.01704 0.00819 0.0055 0.4667 0.7378
-1.000 -0.1147 0.01716 0.00832 0.0063 0.4723 0.7523
-0.750 -0.0872 0.01739 0.00860 0.0070 0.4783 0.7613
-0.500 -0.0621 0.01744 0.00868 0.0077 0.4846 0.7717
-0.250 -0.0373 0.01777 0.00910 0.0091 0.4918 0.7876
0.250 0.0166 0.01812 0.00964 0.0108 0.5087 0.8078
0.500 0.0397 0.01818 0.00977 0.0119 0.5171 0.8200
0.750 0.0714 0.01843 0.01016 0.0121 0.5277 0.8289
1.000 0.0970 0.01852 0.01036 0.0130 0.5362 0.8402
1.250 0.1364 0.01911 0.01108 0.0127 0.5431 0.8598
1.500 0.1704 0.01931 0.01131 0.0123 0.5428 0.8736
1.750 0.2036 0.01934 0.01137 0.0114 0.5406 0.8822
2.000 0.3928 0.02012 0.01241 -0.0157 0.5537 0.9474
2.500 0.4483 0.01957 0.01193 -0.0170 0.5442 0.9559
2.750 0.4770 0.01933 0.01175 -0.0179 0.5383 0.9600
3.000 0.5007 0.01920 0.01164 -0.0177 0.5293 0.9645
3.250 0.5246 0.01908 0.01147 -0.0176 0.5116 0.9682
3.500 0.5543 0.01894 0.01125 -0.0185 0.4851 0.9718
3.750 0.5804 0.01894 0.01117 -0.0189 0.4555 0.9759
4.000 0.6011 0.01913 0.01121 -0.0182 0.4223 0.9796
4.250 0.6223 0.01950 0.01130 -0.0178 0.3946 0.9829
4.500 0.6525 0.01953 0.01148 -0.0190 0.3290 0.9865
4.750 0.6688 0.02194 0.01241 -0.0178 0.3101 0.9892
5.000 0.7079 0.02617 0.01620 -0.0207 0.2749 0.9903
5.250 0.7288 0.02775 0.01814 -0.0205 0.2480 0.9929
5.500 0.7493 0.02709 0.01809 -0.0200 0.2211 0.9970
5.750 0.7670 0.02589 0.01740 -0.0188 0.1938 1.0000
6.000 0.7729 0.02509 0.01699 -0.0157 0.1808 1.0000
6.250 0.7786 0.02488 0.01698 -0.0127 0.1665 1.0000
6.500 0.7832 0.02568 0.01772 -0.0095 0.1497 1.0000
6.750 0.7869 0.02665 0.01865 -0.0061 0.1360 1.0000
7.000 0.7887 0.02761 0.01950 -0.0025 0.1258 1.0000
7.250 0.7971 0.02828 0.02021 0.0004 0.1145 1.0000
7.500 0.8054 0.02928 0.02114 0.0033 0.1066 1.0000
7.750 0.8161 0.03008 0.02193 0.0059 0.0992 1.0000
8.000 0.8294 0.03134 0.02312 0.0079 0.0923 1.0000
8.250 0.8448 0.03246 0.02436 0.0098 0.0865 1.0000
8.500 0.8589 0.03386 0.02570 0.0116 0.0822 1.0000
8.750 0.8787 0.03606 0.02806 0.0127 0.0780 1.0000
9.000 0.8873 0.03747 0.02973 0.0156 0.0747 1.0000
9.250 0.8928 0.03898 0.03142 0.0187 0.0721 1.0000
9.500 0.8961 0.04068 0.03331 0.0221 0.0708 1.0000
9.750 0.8959 0.04241 0.03522 0.0258 0.0696 1.0000
10.000 0.8976 0.04466 0.03768 0.0289 0.0688 1.0000
10.250 0.9023 0.04760 0.04084 0.0308 0.0677 1.0000
10.500 0.9043 0.05142 0.04491 0.0323 0.0669 1.0000
11.000 0.8884 0.05964 0.05372 0.0362 0.0660 1.0000
11.250 0.8751 0.06332 0.05763 0.0383 0.0660 1.0000
11.500 0.8579 0.06711 0.06161 0.0402 0.0661 1.0000
11.750 0.8399 0.07020 0.06494 0.0414 0.0666 1.0000
12.000 0.7187 0.08569 0.08096 0.0358 0.0776 1.0000
12.250 0.7069 0.09190 0.08717 0.0335 0.0791 1.0000
12.500 0.7122 0.09702 0.09228 0.0330 0.0802 1.0000
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Polar data table (+)
Polar graphs
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