GOE 565 AIRFOIL (goe565-il) Xfoil prediction polar at RE=500,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: GOE 565 AIRFOIL (goe565-il) Reynolds number: 500,000 Max Cl/Cd: 98.73 at α=2.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe565-il-500000.txt Download as CSV file: xf-goe565-il-500000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 565 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.500 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.250 -0.4422 0.08955 0.08721 -0.0311 1.0000 0.0269
-9.000 -0.4435 0.08604 0.08373 -0.0322 1.0000 0.0277
-8.750 -0.7004 0.03033 0.02709 -0.0550 1.0000 0.0168
-8.500 -0.6998 0.02577 0.02197 -0.0533 1.0000 0.0174
-8.250 -0.6722 0.02312 0.01886 -0.0551 0.9979 0.0185
-8.000 -0.6448 0.02042 0.01574 -0.0567 0.9957 0.0198
-7.750 -0.6114 0.02011 0.01540 -0.0583 0.9941 0.0207
-7.500 -0.5815 0.01946 0.01463 -0.0591 0.9917 0.0220
-7.250 -0.5511 0.01836 0.01327 -0.0601 0.9892 0.0235
-7.000 -0.5206 0.01687 0.01154 -0.0612 0.9869 0.0252
-6.750 -0.4854 0.01706 0.01177 -0.0629 0.9851 0.0269
-6.500 -0.4502 0.01688 0.01150 -0.0646 0.9837 0.0292
-6.250 -0.4216 0.01591 0.01033 -0.0650 0.9808 0.0315
-6.000 -0.3911 0.01557 0.01001 -0.0658 0.9776 0.0336
-5.750 -0.3572 0.01551 0.00990 -0.0672 0.9751 0.0363
-5.500 -0.3217 0.01548 0.00976 -0.0687 0.9732 0.0386
-5.250 -0.2893 0.01398 0.00816 -0.0701 0.9717 0.0417
-5.000 -0.2533 0.01368 0.00783 -0.0720 0.9705 0.0445
-4.750 -0.2256 0.01333 0.00743 -0.0719 0.9659 0.0469
-4.500 -0.1919 0.01308 0.00711 -0.0731 0.9628 0.0489
-4.250 -0.1570 0.01205 0.00599 -0.0747 0.9605 0.0512
-4.000 -0.1194 0.01133 0.00527 -0.0768 0.9587 0.0543
-3.750 -0.0816 0.01082 0.00474 -0.0788 0.9569 0.0567
-3.500 -0.0523 0.01043 0.00432 -0.0790 0.9514 0.0586
-3.250 -0.0209 0.01010 0.00395 -0.0796 0.9468 0.0608
-3.000 0.0122 0.00971 0.00353 -0.0806 0.9436 0.0625
-2.750 0.0415 0.00929 0.00310 -0.0807 0.9388 0.0662
-2.500 0.0690 0.00903 0.00285 -0.0805 0.9323 0.0701
-2.000 0.1259 0.00854 0.00238 -0.0802 0.9179 0.0861
-1.750 0.1534 0.00820 0.00213 -0.0798 0.9087 0.1175
-1.500 0.1783 0.00757 0.00194 -0.0792 0.8982 0.2485
-1.250 0.2035 0.00731 0.00185 -0.0785 0.8874 0.3134
-1.000 0.2298 0.00706 0.00176 -0.0781 0.8779 0.3712
-0.750 0.2551 0.00669 0.00170 -0.0775 0.8679 0.4748
-0.500 0.2785 0.00634 0.00171 -0.0764 0.8561 0.5943
-0.250 0.3024 0.00607 0.00169 -0.0753 0.8445 0.6838
0.000 0.3245 0.00570 0.00167 -0.0738 0.8330 0.7819
0.250 0.3873 0.00537 0.00166 -0.0808 0.8239 0.9792
0.500 0.4311 0.00539 0.00162 -0.0841 0.8103 1.0000
0.750 0.4554 0.00545 0.00158 -0.0832 0.7929 1.0000
1.000 0.4797 0.00552 0.00157 -0.0822 0.7748 1.0000
1.250 0.5043 0.00561 0.00158 -0.0813 0.7581 1.0000
1.500 0.5286 0.00572 0.00160 -0.0803 0.7397 1.0000
1.750 0.5526 0.00585 0.00163 -0.0793 0.7198 1.0000
2.000 0.5768 0.00597 0.00168 -0.0784 0.6995 1.0000
2.250 0.6004 0.00612 0.00175 -0.0773 0.6761 1.0000
2.500 0.6230 0.00631 0.00180 -0.0760 0.6384 1.0000
2.750 0.6434 0.00662 0.00187 -0.0743 0.5826 1.0000
3.000 0.6641 0.00699 0.00200 -0.0728 0.5332 1.0000
3.250 0.6861 0.00732 0.00216 -0.0716 0.4946 1.0000
3.500 0.7084 0.00766 0.00233 -0.0704 0.4548 1.0000
3.750 0.7310 0.00800 0.00251 -0.0694 0.4144 1.0000
4.000 0.7534 0.00837 0.00270 -0.0683 0.3731 1.0000
4.250 0.7755 0.00879 0.00294 -0.0672 0.3318 1.0000
4.500 0.7975 0.00924 0.00319 -0.0661 0.2875 1.0000
4.750 0.8198 0.00968 0.00345 -0.0652 0.2459 1.0000
5.000 0.8418 0.01017 0.00374 -0.0642 0.2055 1.0000
5.250 0.8642 0.01062 0.00406 -0.0632 0.1795 1.0000
5.500 0.8873 0.01102 0.00438 -0.0624 0.1597 1.0000
5.750 0.9101 0.01145 0.00470 -0.0615 0.1345 1.0000
6.000 0.9283 0.01236 0.00522 -0.0599 0.0658 1.0000
6.250 0.9489 0.01304 0.00577 -0.0587 0.0465 1.0000
6.500 0.9713 0.01353 0.00627 -0.0577 0.0419 1.0000
6.750 0.9917 0.01424 0.00700 -0.0563 0.0372 1.0000
7.000 1.0141 0.01469 0.00752 -0.0553 0.0354 1.0000
7.250 1.0353 0.01525 0.00815 -0.0542 0.0330 1.0000
7.500 1.0544 0.01602 0.00895 -0.0527 0.0304 1.0000
7.750 1.0701 0.01712 0.01013 -0.0507 0.0282 1.0000
8.000 1.0912 0.01766 0.01074 -0.0496 0.0269 1.0000
8.250 1.1110 0.01831 0.01144 -0.0482 0.0252 1.0000
8.500 1.1299 0.01901 0.01218 -0.0469 0.0236 1.0000
8.750 1.1413 0.02060 0.01382 -0.0445 0.0219 1.0000
9.000 1.1601 0.02134 0.01465 -0.0430 0.0210 1.0000
9.250 1.1786 0.02208 0.01548 -0.0416 0.0201 1.0000
9.500 1.1959 0.02294 0.01642 -0.0400 0.0192 1.0000
9.750 1.2127 0.02374 0.01730 -0.0384 0.0183 1.0000
10.000 1.2283 0.02456 0.01815 -0.0368 0.0175 1.0000
10.250 1.2387 0.02658 0.02022 -0.0346 0.0165 1.0000
10.500 1.2514 0.02759 0.02138 -0.0324 0.0160 1.0000
10.750 1.2634 0.02864 0.02256 -0.0302 0.0156 1.0000
11.000 1.2742 0.02977 0.02383 -0.0279 0.0150 1.0000
11.250 1.2837 0.03105 0.02524 -0.0256 0.0145 1.0000
11.500 1.2918 0.03245 0.02678 -0.0233 0.0141 1.0000
11.750 1.2983 0.03383 0.02828 -0.0210 0.0137 1.0000
12.000 1.3040 0.03507 0.02962 -0.0188 0.0134 1.0000
12.250 1.3081 0.03670 0.03135 -0.0166 0.0131 1.0000
12.500 1.3105 0.03861 0.03337 -0.0146 0.0129 1.0000
12.750 1.3076 0.04144 0.03636 -0.0126 0.0126 1.0000
13.250 1.2864 0.04857 0.04398 -0.0090 0.0122 1.0000
13.500 1.2778 0.05171 0.04732 -0.0081 0.0121 1.0000
13.750 1.2686 0.05508 0.05090 -0.0078 0.0120 1.0000
14.000 1.2549 0.05944 0.05548 -0.0083 0.0118 1.0000
14.250 1.2387 0.06458 0.06082 -0.0096 0.0118 1.0000
14.500 1.2217 0.07026 0.06671 -0.0118 0.0117 1.0000
14.750 1.1996 0.07737 0.07401 -0.0151 0.0119 1.0000
15.000 1.1818 0.08435 0.08118 -0.0192 0.0117 1.0000
15.250 1.1598 0.09278 0.08978 -0.0244 0.0118 1.0000
15.500 1.1327 0.10293 0.10010 -0.0307 0.0120 1.0000
15.750 1.1084 0.11319 0.11051 -0.0372 0.0121 1.0000
16.000 1.0797 0.12510 0.12255 -0.0447 0.0123 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 565 AIRFOIL (goe565-il)