DAE-51 AIRFOIL (dae51-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
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Airfoil: DAE-51 AIRFOIL (dae51-il) Reynolds number: 100,000 Max Cl/Cd: 62.55 at α=7° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-dae51-il-100000.txt Download as CSV file: xf-dae51-il-100000.csv |
XFOIL Version 6.96
Calculated polar for: DAE-51 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
------ -------- --------- --------- -------- -------- --------
-8.500 -0.3478 0.10365 0.09893 -0.0389 1.0000 0.0812
-8.250 -0.3653 0.10299 0.09843 -0.0397 1.0000 0.0816
-8.000 -0.3845 0.10228 0.09785 -0.0393 1.0000 0.0818
-7.750 -0.3435 0.09416 0.08963 -0.0349 1.0000 0.0851
-7.500 -0.3524 0.09278 0.08834 -0.0318 1.0000 0.0866
-7.250 -0.3694 0.09202 0.08767 -0.0279 1.0000 0.0875
-7.000 -0.3844 0.09094 0.08669 -0.0251 1.0000 0.0887
-6.750 -0.3966 0.08947 0.08529 -0.0235 1.0000 0.0904
-6.500 -0.4074 0.08769 0.08359 -0.0234 1.0000 0.0925
-6.250 -0.4178 0.08537 0.08128 -0.0331 1.0000 0.0956
-6.000 -0.4181 0.08118 0.07710 -0.0354 1.0000 0.0969
-5.750 -0.4171 0.07879 0.07478 -0.0294 1.0000 0.0986
-5.500 -0.4142 0.07663 0.07265 -0.0270 1.0000 0.1014
-5.250 -0.3822 0.07044 0.06624 -0.0423 0.9963 0.1119
-5.000 -0.3637 0.06799 0.06388 -0.0394 0.9922 0.1176
-4.750 -0.3268 0.06284 0.05859 -0.0486 0.9863 0.1298
-4.500 -0.2903 0.05832 0.05391 -0.0563 0.9807 0.1436
-4.250 -0.2550 0.05448 0.04992 -0.0620 0.9748 0.1587
-4.000 -0.2182 0.05111 0.04646 -0.0666 0.9702 0.1759
-3.750 -0.1842 0.04802 0.04317 -0.0717 0.9631 0.2026
-3.500 -0.0968 0.03431 0.02752 -0.0867 0.9619 0.0856
-3.250 -0.0474 0.03101 0.02344 -0.0908 0.9585 0.0791
-3.000 -0.0070 0.02850 0.02043 -0.0934 0.9535 0.0770
-2.750 0.0337 0.02653 0.01799 -0.0956 0.9481 0.0766
-2.500 0.0797 0.02535 0.01634 -0.0986 0.9438 0.0818
-2.250 0.1152 0.02389 0.01492 -0.1003 0.9377 0.0892
-2.000 0.1547 0.02282 0.01385 -0.1025 0.9320 0.1023
-1.750 0.2013 0.02170 0.01288 -0.1060 0.9281 0.1379
-1.500 0.2351 0.01921 0.01250 -0.1077 0.9228 0.5874
-1.250 0.2602 0.01835 0.01232 -0.1052 0.9157 1.0000
-1.000 0.2947 0.01860 0.01224 -0.1067 0.9075 1.0000
-0.750 0.3352 0.01874 0.01208 -0.1092 0.9006 1.0000
-0.500 0.3671 0.01900 0.01213 -0.1101 0.8920 1.0000
-0.250 0.4057 0.01909 0.01203 -0.1120 0.8852 1.0000
0.000 0.4338 0.01937 0.01217 -0.1122 0.8755 1.0000
0.250 0.4726 0.01936 0.01202 -0.1138 0.8694 1.0000
0.500 0.4980 0.01967 0.01223 -0.1135 0.8588 1.0000
0.750 0.5365 0.01956 0.01202 -0.1148 0.8534 1.0000
1.000 0.5603 0.01988 0.01227 -0.1141 0.8420 1.0000
1.250 0.5886 0.02005 0.01238 -0.1140 0.8328 1.0000
1.500 0.6211 0.01997 0.01226 -0.1142 0.8253 1.0000
1.750 0.6461 0.02022 0.01247 -0.1136 0.8144 1.0000
2.000 0.6804 0.01995 0.01217 -0.1137 0.8087 1.0000
2.250 0.7045 0.02019 0.01241 -0.1129 0.7968 1.0000
2.500 0.7304 0.02032 0.01255 -0.1122 0.7862 1.0000
2.750 0.7627 0.02000 0.01220 -0.1119 0.7794 1.0000
3.000 0.7873 0.02016 0.01239 -0.1110 0.7673 1.0000
3.250 0.8133 0.02021 0.01248 -0.1102 0.7561 1.0000
3.500 0.8432 0.01995 0.01223 -0.1096 0.7475 1.0000
3.750 0.8705 0.01983 0.01214 -0.1088 0.7364 1.0000
4.000 0.8962 0.01984 0.01222 -0.1078 0.7238 1.0000
4.250 0.9227 0.01977 0.01220 -0.1069 0.7114 1.0000
4.500 0.9499 0.01964 0.01211 -0.1060 0.6990 1.0000
4.750 0.9774 0.01946 0.01197 -0.1052 0.6862 1.0000
5.000 1.0046 0.01930 0.01189 -0.1042 0.6723 1.0000
5.250 1.0310 0.01920 0.01183 -0.1032 0.6570 1.0000
5.500 1.0571 0.01911 0.01180 -0.1022 0.6405 1.0000
5.750 1.0833 0.01901 0.01177 -0.1012 0.6228 1.0000
6.000 1.1100 0.01887 0.01163 -0.1001 0.6041 1.0000
6.250 1.1337 0.01893 0.01176 -0.0988 0.5816 1.0000
6.500 1.1588 0.01890 0.01170 -0.0975 0.5582 1.0000
6.750 1.1811 0.01902 0.01186 -0.0960 0.5304 1.0000
7.000 1.2023 0.01922 0.01204 -0.0944 0.5000 1.0000
7.250 1.2214 0.01955 0.01240 -0.0926 0.4655 1.0000
7.500 1.2394 0.01999 0.01277 -0.0906 0.4281 1.0000
7.750 1.2547 0.02062 0.01333 -0.0885 0.3850 1.0000
8.000 1.2673 0.02149 0.01404 -0.0861 0.3366 1.0000
8.250 1.2758 0.02270 0.01499 -0.0833 0.2820 1.0000
8.500 1.2794 0.02438 0.01630 -0.0801 0.2207 1.0000
8.750 1.2779 0.02659 0.01803 -0.0764 0.1664 1.0000
9.000 1.2764 0.02882 0.01994 -0.0726 0.1304 1.0000
9.250 1.2769 0.03095 0.02188 -0.0691 0.1098 1.0000
9.500 1.2817 0.03295 0.02382 -0.0663 0.0950 1.0000
9.750 1.2903 0.03499 0.02578 -0.0640 0.0846 1.0000
10.000 1.3058 0.03716 0.02779 -0.0625 0.0762 1.0000
10.250 1.3212 0.03898 0.02982 -0.0608 0.0694 1.0000
10.500 1.3520 0.04182 0.03251 -0.0611 0.0632 1.0000
10.750 1.3711 0.04407 0.03511 -0.0598 0.0596 1.0000
11.000 1.3864 0.04626 0.03747 -0.0585 0.0559 1.0000
11.250 1.2932 0.04006 0.03200 -0.0437 0.0571 1.0000
11.500 1.3152 0.04366 0.03560 -0.0435 0.0536 1.0000
11.750 1.4233 0.05683 0.04895 -0.0542 0.0510 1.0000
12.000 1.4175 0.06003 0.05254 -0.0511 0.0505 1.0000
12.250 1.4076 0.06336 0.05622 -0.0480 0.0501 1.0000
12.500 1.3942 0.06690 0.06010 -0.0454 0.0497 1.0000
12.750 1.3783 0.07075 0.06427 -0.0432 0.0495 1.0000
13.000 1.3601 0.07501 0.06882 -0.0417 0.0493 1.0000
13.250 1.3395 0.07970 0.07380 -0.0408 0.0494 1.0000
13.500 1.3171 0.08494 0.07930 -0.0408 0.0496 1.0000
13.750 1.2932 0.09072 0.08532 -0.0417 0.0499 1.0000
14.000 1.2685 0.09711 0.09192 -0.0434 0.0504 1.0000
14.250 1.2432 0.10410 0.09911 -0.0462 0.0509 1.0000
14.500 1.2190 0.11164 0.10680 -0.0497 0.0515 1.0000
14.750 1.1965 0.11967 0.11496 -0.0538 0.0520 1.0000
15.000 1.1790 0.12778 0.12313 -0.0576 0.0526 1.0000
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Polar data table (+)
Polar graphs
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