GOE 562 AIRFOIL (goe562-il) Xfoil prediction polar at RE=100,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: GOE 562 AIRFOIL (goe562-il) Reynolds number: 100,000 Max Cl/Cd: 49.98 at α=6° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe562-il-100000-n5.txt Download as CSV file: xf-goe562-il-100000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 562 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
------ -------- --------- --------- -------- -------- --------
-9.750 -0.1635 0.11114 0.10681 -0.0528 0.9443 0.0473
-9.500 -0.1586 0.10795 0.10362 -0.0599 0.9199 0.0495
-9.250 -0.1472 0.10384 0.09948 -0.0692 0.9031 0.0500
-9.000 -0.1173 0.09738 0.09296 -0.0714 0.8911 0.0513
-8.750 -0.0831 0.09239 0.08786 -0.0765 0.8763 0.0535
-8.500 -0.0509 0.08741 0.08272 -0.0848 0.8585 0.0562
-8.250 -0.0225 0.08239 0.07755 -0.0954 0.8383 0.0595
-8.000 -0.0233 0.07852 0.07354 -0.1076 0.8107 0.0615
-7.500 0.0220 0.07045 0.06523 -0.1110 0.7743 0.0658
-7.250 0.0319 0.06756 0.06222 -0.1130 0.7574 0.0680
-7.000 0.0390 0.06452 0.05906 -0.1153 0.7424 0.0701
-6.750 0.0317 0.06195 0.05616 -0.1194 0.7296 0.0752
-6.500 0.0330 0.05953 0.05341 -0.1194 0.7180 0.0756
-6.250 0.0525 0.05504 0.04910 -0.1192 0.7078 0.0778
-5.500 0.0719 0.04271 0.03547 -0.1151 0.6834 0.0501
-5.250 0.0833 0.04007 0.03270 -0.1137 0.6755 0.0510
-5.000 0.1017 0.03905 0.03172 -0.1132 0.6673 0.0538
-4.750 0.1148 0.03693 0.02929 -0.1113 0.6602 0.0540
-4.500 0.1270 0.03451 0.02644 -0.1088 0.6533 0.0529
-4.250 0.1429 0.03241 0.02384 -0.1066 0.6475 0.0521
-4.000 0.1587 0.03090 0.02205 -0.1045 0.6405 0.0524
-3.750 0.1785 0.02981 0.02073 -0.1032 0.6344 0.0541
-3.500 0.1980 0.02867 0.01930 -0.1017 0.6283 0.0553
-3.250 0.2177 0.02749 0.01781 -0.1000 0.6220 0.0556
-3.000 0.2409 0.02638 0.01633 -0.0989 0.6167 0.0557
-2.750 0.2626 0.02545 0.01514 -0.0975 0.6107 0.0561
-2.500 0.2863 0.02463 0.01408 -0.0965 0.6051 0.0565
-2.250 0.3126 0.02399 0.01317 -0.0961 0.6002 0.0584
-2.000 0.3359 0.02348 0.01249 -0.0950 0.5943 0.0600
-1.750 0.3611 0.02296 0.01179 -0.0943 0.5888 0.0609
-1.500 0.3890 0.02245 0.01109 -0.0942 0.5842 0.0616
-1.250 0.4132 0.02194 0.01055 -0.0935 0.5790 0.0624
-1.000 0.4372 0.02144 0.01006 -0.0928 0.5739 0.0636
-0.750 0.4621 0.02106 0.00964 -0.0922 0.5695 0.0654
-0.500 0.4854 0.02084 0.00939 -0.0913 0.5649 0.0690
-0.250 0.5068 0.02071 0.00923 -0.0900 0.5596 0.0727
0.000 0.5313 0.02059 0.00900 -0.0893 0.5551 0.0753
0.250 0.5591 0.02041 0.00871 -0.0893 0.5516 0.0795
0.500 0.5827 0.02040 0.00870 -0.0886 0.5465 0.0853
0.750 0.6111 0.02033 0.00866 -0.0888 0.5416 0.1025
1.000 0.7993 0.01867 0.00887 -0.1239 0.5320 1.0000
1.250 0.8190 0.01888 0.00899 -0.1223 0.5263 1.0000
1.500 0.8414 0.01901 0.00894 -0.1212 0.5210 1.0000
1.750 0.8596 0.01922 0.00910 -0.1194 0.5147 1.0000
2.000 0.8792 0.01938 0.00916 -0.1178 0.5082 1.0000
2.250 0.9004 0.01952 0.00917 -0.1165 0.5026 1.0000
2.500 0.9178 0.01976 0.00940 -0.1146 0.4962 1.0000
2.750 0.9388 0.01995 0.00951 -0.1133 0.4913 1.0000
3.000 0.9602 0.02017 0.00964 -0.1121 0.4871 1.0000
3.250 0.9778 0.02046 0.00997 -0.1103 0.4819 1.0000
3.500 0.9976 0.02069 0.01015 -0.1088 0.4769 1.0000
3.750 1.0201 0.02088 0.01025 -0.1079 0.4730 1.0000
4.000 1.0361 0.02122 0.01065 -0.1058 0.4676 1.0000
4.250 1.0544 0.02148 0.01092 -0.1041 0.4623 1.0000
4.500 1.0756 0.02169 0.01106 -0.1029 0.4581 1.0000
4.750 1.0916 0.02203 0.01145 -0.1009 0.4528 1.0000
5.000 1.1085 0.02234 0.01180 -0.0990 0.4475 1.0000
5.250 1.1284 0.02258 0.01200 -0.0976 0.4431 1.0000
5.500 1.1443 0.02293 0.01240 -0.0955 0.4381 1.0000
5.750 1.1592 0.02330 0.01282 -0.0933 0.4326 1.0000
6.000 1.1775 0.02356 0.01309 -0.0917 0.4279 1.0000
6.250 1.1921 0.02395 0.01352 -0.0895 0.4226 1.0000
6.500 1.2052 0.02437 0.01402 -0.0870 0.4170 1.0000
6.750 1.2221 0.02467 0.01432 -0.0852 0.4122 1.0000
7.000 1.2334 0.02512 0.01484 -0.0824 0.4066 1.0000
7.250 1.2428 0.02554 0.01534 -0.0793 0.4009 1.0000
7.500 1.2577 0.02583 0.01562 -0.0771 0.3961 1.0000
7.750 1.2636 0.02639 0.01629 -0.0735 0.3901 1.0000
8.000 1.2727 0.02687 0.01685 -0.0706 0.3844 1.0000
8.250 1.2871 0.02724 0.01720 -0.0685 0.3796 1.0000
8.500 1.2904 0.02799 0.01809 -0.0648 0.3731 1.0000
8.750 1.3008 0.02852 0.01867 -0.0623 0.3678 1.0000
9.000 1.3106 0.02916 0.01938 -0.0599 0.3626 1.0000
9.250 1.3152 0.03002 0.02036 -0.0568 0.3565 1.0000
9.500 1.3274 0.03062 0.02098 -0.0549 0.3520 1.0000
9.750 1.3334 0.03157 0.02204 -0.0523 0.3470 1.0000
10.000 1.3385 0.03262 0.02321 -0.0498 0.3420 1.0000
10.250 1.3498 0.03340 0.02404 -0.0481 0.3380 1.0000
10.500 1.3567 0.03446 0.02520 -0.0460 0.3335 1.0000
10.750 1.3564 0.03595 0.02683 -0.0434 0.3282 1.0000
11.000 1.3649 0.03672 0.02758 -0.0416 0.3217 1.0000
11.250 1.3557 0.03898 0.03001 -0.0389 0.3155 1.0000
11.500 1.3575 0.04036 0.03138 -0.0371 0.3081 1.0000
11.750 1.3504 0.04281 0.03398 -0.0351 0.3022 1.0000
12.000 1.3476 0.04492 0.03616 -0.0336 0.2958 1.0000
12.250 1.3448 0.04716 0.03849 -0.0322 0.2896 1.0000
12.500 1.3363 0.05014 0.04161 -0.0310 0.2837 1.0000
12.750 1.3397 0.05186 0.04332 -0.0300 0.2771 1.0000
13.000 1.3257 0.05575 0.04739 -0.0292 0.2717 1.0000
13.250 1.3219 0.05842 0.05012 -0.0284 0.2651 1.0000
13.500 1.3112 0.06210 0.05390 -0.0279 0.2585 1.0000
13.750 1.2998 0.06603 0.05795 -0.0276 0.2525 1.0000
14.000 1.3030 0.06804 0.05998 -0.0270 0.2462 1.0000
14.250 1.2856 0.07308 0.06523 -0.0272 0.2419 1.0000
14.500 1.2739 0.07736 0.06962 -0.0273 0.2353 1.0000
14.750 1.2660 0.08116 0.07350 -0.0273 0.2281 1.0000
15.000 1.2467 0.08688 0.07939 -0.0280 0.2225 1.0000
15.250 1.2354 0.09136 0.08393 -0.0285 0.2123 1.0000
15.500 1.2174 0.09721 0.08997 -0.0294 0.2086 1.0000
15.750 1.2025 0.10255 0.09543 -0.0304 0.1987 1.0000
16.000 1.1771 0.10991 0.10301 -0.0320 0.1962 1.0000
16.250 1.1482 0.11808 0.11134 -0.0340 0.1935 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 562 AIRFOIL (goe562-il)