GOE 527 AIRFOIL (goe527-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: GOE 527 AIRFOIL (goe527-il) Reynolds number: 100,000 Max Cl/Cd: 36.4 at α=8.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe527-il-100000.txt Download as CSV file: xf-goe527-il-100000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 527 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
------ -------- --------- --------- -------- -------- --------
-9.250 -0.1010 0.09953 0.09491 -0.0819 0.9431 0.1375
-9.000 -0.0584 0.09541 0.09074 -0.0841 0.9400 0.1425
-8.750 -0.0823 0.09285 0.08821 -0.0962 0.9256 0.1513
-8.500 -0.0327 0.08708 0.08239 -0.0965 0.9244 0.1542
-8.250 0.0076 0.08320 0.07846 -0.1000 0.9213 0.1600
-8.000 -0.0682 0.08163 0.07691 -0.1109 0.8928 0.1682
-7.750 0.0342 0.07630 0.07154 -0.1075 0.9009 0.1738
-7.500 0.0128 0.07401 0.06923 -0.1130 0.8847 0.1842
-7.250 0.0248 0.07080 0.06603 -0.1118 0.8730 0.1883
-7.000 0.0623 0.06794 0.06310 -0.1131 0.8677 0.1965
-6.750 0.0236 0.06552 0.06066 -0.1150 0.8506 0.2054
-6.500 0.0496 0.06370 0.05885 -0.1123 0.8417 0.2096
-6.250 -0.0549 0.04648 0.04029 -0.1196 0.8227 0.1223
-6.000 -0.0361 0.04339 0.03701 -0.1193 0.8152 0.1179
-5.750 -0.0243 0.04054 0.03381 -0.1181 0.8073 0.1167
-5.500 -0.0137 0.03831 0.03120 -0.1162 0.7980 0.1168
-5.250 0.0121 0.03559 0.02798 -0.1163 0.7933 0.1175
-5.000 0.0180 0.03457 0.02668 -0.1131 0.7822 0.1181
-4.750 0.0446 0.03284 0.02439 -0.1127 0.7768 0.1205
-4.500 0.0623 0.03156 0.02304 -0.1113 0.7689 0.1242
-4.250 0.0848 0.03075 0.02215 -0.1103 0.7613 0.1289
-4.000 0.1174 0.02948 0.02046 -0.1105 0.7569 0.1348
-3.750 0.1297 0.02903 0.02004 -0.1083 0.7480 0.1400
-3.500 0.1566 0.02838 0.01924 -0.1078 0.7419 0.1486
-3.250 0.1912 0.02728 0.01809 -0.1085 0.7379 0.1590
-3.000 0.2010 0.02757 0.01829 -0.1057 0.7286 0.1670
-2.750 0.2285 0.02697 0.01776 -0.1055 0.7232 0.1785
-2.500 0.2627 0.02625 0.01701 -0.1061 0.7195 0.1927
-2.250 0.2706 0.02679 0.01752 -0.1030 0.7104 0.2023
-2.000 0.2976 0.02635 0.01718 -0.1027 0.7050 0.2171
-1.750 0.3311 0.02573 0.01660 -0.1031 0.7014 0.2367
-1.500 0.3373 0.02631 0.01733 -0.1000 0.6933 0.2515
-1.250 0.3590 0.02608 0.01728 -0.0988 0.6875 0.2834
-1.000 0.3872 0.02491 0.01689 -0.0984 0.6838 0.4238
-0.750 0.3991 0.02438 0.01742 -0.0948 0.6786 0.6812
-0.500 0.5575 0.02383 0.01710 -0.1175 0.6732 1.0000
-0.250 0.5753 0.02405 0.01713 -0.1159 0.6679 1.0000
0.000 0.6042 0.02401 0.01684 -0.1158 0.6642 1.0000
0.250 0.5932 0.02544 0.01827 -0.1102 0.6559 1.0000
0.500 0.6114 0.02580 0.01850 -0.1086 0.6504 1.0000
0.750 0.6448 0.02565 0.01815 -0.1090 0.6468 1.0000
1.000 0.6353 0.02718 0.01968 -0.1037 0.6387 1.0000
1.250 0.6497 0.02775 0.02016 -0.1015 0.6327 1.0000
1.500 0.6877 0.02742 0.01966 -0.1025 0.6292 1.0000
1.750 0.6792 0.02900 0.02124 -0.0974 0.6211 1.0000
2.000 0.6936 0.02957 0.02174 -0.0953 0.6145 1.0000
2.250 0.7420 0.02871 0.02069 -0.0975 0.6112 1.0000
2.500 0.7266 0.03055 0.02256 -0.0916 0.6015 1.0000
2.750 0.7608 0.03017 0.02208 -0.0919 0.5961 1.0000
3.000 0.8135 0.02908 0.02081 -0.0947 0.5927 1.0000
3.250 0.7888 0.03135 0.02316 -0.0876 0.5824 1.0000
3.500 0.8272 0.03087 0.02258 -0.0885 0.5778 1.0000
3.750 0.8778 0.02996 0.02152 -0.0911 0.5747 1.0000
4.000 0.8437 0.03268 0.02435 -0.0830 0.5636 1.0000
4.250 0.8878 0.03197 0.02356 -0.0847 0.5597 1.0000
4.500 0.9397 0.03103 0.02249 -0.0874 0.5567 1.0000
4.750 0.8940 0.03427 0.02586 -0.0780 0.5448 1.0000
5.000 0.9479 0.03310 0.02460 -0.0808 0.5414 1.0000
5.250 1.0083 0.03176 0.02312 -0.0846 0.5384 1.0000
5.500 0.9548 0.03514 0.02665 -0.0742 0.5261 1.0000
5.750 1.0194 0.03355 0.02497 -0.0783 0.5229 1.0000
6.000 1.0847 0.03213 0.02340 -0.0829 0.5199 1.0000
6.250 1.0231 0.03563 0.02709 -0.0712 0.5075 1.0000
6.500 1.0923 0.03398 0.02535 -0.0762 0.5042 1.0000
6.750 1.0238 0.03806 0.02955 -0.0644 0.4917 1.0000
7.000 1.0929 0.03605 0.02748 -0.0686 0.4885 1.0000
7.250 1.1780 0.03388 0.02518 -0.0757 0.4852 1.0000
7.500 1.0994 0.03804 0.02950 -0.0624 0.4727 1.0000
7.750 1.1841 0.03546 0.02682 -0.0686 0.4693 1.0000
8.000 1.1098 0.04012 0.03163 -0.0573 0.4570 1.0000
8.250 1.1789 0.03783 0.02927 -0.0609 0.4535 1.0000
8.500 1.2789 0.03513 0.02642 -0.0696 0.4500 1.0000
8.750 1.1674 0.04122 0.03276 -0.0541 0.4374 1.0000
9.000 1.2674 0.03738 0.02879 -0.0607 0.4345 1.0000
9.250 1.1450 0.04611 0.03774 -0.0480 0.4203 1.0000
9.500 1.2605 0.03997 0.03147 -0.0538 0.4190 1.0000
9.750 1.0618 0.05843 0.05017 -0.0425 0.3969 1.0000
10.000 1.1265 0.05389 0.04561 -0.0424 0.3963 1.0000
10.250 1.2120 0.04807 0.03975 -0.0438 0.3950 1.0000
10.500 1.1192 0.05930 0.05109 -0.0400 0.3807 1.0000
10.750 1.1937 0.05362 0.04538 -0.0401 0.3793 1.0000
11.000 1.1101 0.06528 0.05713 -0.0381 0.3650 1.0000
11.250 1.1752 0.06017 0.05203 -0.0375 0.3643 1.0000
11.500 1.2575 0.05421 0.04604 -0.0380 0.3638 1.0000
11.750 1.1472 0.06904 0.06100 -0.0362 0.3498 1.0000
12.000 1.2134 0.06384 0.05580 -0.0355 0.3495 1.0000
12.250 1.3079 0.05650 0.04843 -0.0360 0.3490 1.0000
13.000 1.3908 0.05572 0.04772 -0.0349 0.3343 1.0000
13.250 0.9245 0.12317 0.11550 -0.0451 0.2874 1.0000
13.500 0.8771 0.13447 0.12685 -0.0489 0.2757 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 527 AIRFOIL (goe527-il)