GOE 559 AIRFOIL (goe559-il) Xfoil prediction polar at RE=500,000 Ncrit=5
| Details | Polar file |
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Airfoil: GOE 559 AIRFOIL (goe559-il) Reynolds number: 500,000 Max Cl/Cd: 64.19 at α=1.5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe559-il-500000-n5.txt Download as CSV file: xf-goe559-il-500000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 559 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 = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.500 -0.3342 0.10583 0.10372 -0.0340 1.0000 0.0015
-9.250 -0.3299 0.10259 0.10050 -0.0351 1.0000 0.0017
-9.000 -0.3272 0.09933 0.09727 -0.0360 0.9947 0.0017
-8.750 -0.3101 0.09447 0.09240 -0.0417 0.9681 0.0017
-8.500 -0.2687 0.08710 0.08494 -0.0556 0.9333 0.0017
-8.250 -0.2301 0.08043 0.07806 -0.0690 0.8800 0.0017
-8.000 -0.2277 0.07727 0.07455 -0.0707 0.7892 0.0017
-7.750 -0.2289 0.07473 0.07179 -0.0709 0.7418 0.0017
-7.500 -0.2334 0.07126 0.06825 -0.0732 0.7190 0.0017
-7.250 -0.2364 0.06801 0.06493 -0.0752 0.7041 0.0017
-7.000 -0.2339 0.06457 0.06140 -0.0769 0.6904 0.0017
-6.750 -0.2290 0.06090 0.05762 -0.0786 0.6788 0.0017
-6.500 -0.2230 0.05774 0.05434 -0.0793 0.6668 0.0017
-6.250 -0.2128 0.05425 0.05068 -0.0799 0.6554 0.0018
-6.000 -0.2027 0.05121 0.04749 -0.0799 0.6432 0.0018
-5.750 -0.1907 0.04817 0.04426 -0.0796 0.6302 0.0018
-5.500 -0.1767 0.04514 0.04102 -0.0789 0.6170 0.0018
-5.250 -0.1622 0.04233 0.03799 -0.0782 0.6027 0.0018
-5.000 -0.1531 0.03846 0.03393 -0.0776 0.5900 0.0019
-4.750 -0.1393 0.03550 0.03074 -0.0768 0.5775 0.0020
-4.500 -0.1221 0.03289 0.02788 -0.0758 0.5664 0.0021
-4.250 -0.1028 0.03044 0.02515 -0.0745 0.5567 0.0024
-4.000 -0.0819 0.02826 0.02270 -0.0732 0.5480 0.0025
-3.750 -0.0502 0.02824 0.02231 -0.0711 0.5399 0.0037
-3.500 -0.0353 0.02437 0.01823 -0.0702 0.5333 0.0040
-3.000 0.0104 0.02081 0.01412 -0.0680 0.5207 0.0049
-2.750 0.0356 0.01949 0.01255 -0.0669 0.5146 0.0055
-2.500 0.0637 0.01921 0.01201 -0.0659 0.5088 0.0064
-2.250 0.0896 0.01863 0.01112 -0.0650 0.5032 0.0066
-1.750 0.1390 0.01461 0.00684 -0.0639 0.4939 0.0087
-1.500 0.1655 0.01378 0.00588 -0.0634 0.4893 0.0106
-1.250 0.1928 0.01399 0.00585 -0.0629 0.4850 0.0128
-1.000 0.2160 0.01217 0.00411 -0.0619 0.4816 0.0087
-0.750 0.2416 0.01171 0.00346 -0.0611 0.4777 0.0037
-0.500 0.2646 0.01104 0.00273 -0.0602 0.4739 0.0032
-0.250 0.2889 0.01062 0.00220 -0.0595 0.4703 0.0029
0.000 0.3145 0.01036 0.00181 -0.0592 0.4668 0.0029
0.250 0.3408 0.01023 0.00154 -0.0590 0.4630 0.0033
0.500 0.3782 0.00765 0.00166 -0.0618 0.4593 0.9593
0.750 0.4133 0.00775 0.00165 -0.0638 0.4560 0.9672
1.000 0.4468 0.00782 0.00169 -0.0654 0.4530 0.9736
1.250 0.4784 0.00790 0.00175 -0.0666 0.4499 0.9801
1.500 0.5122 0.00798 0.00185 -0.0683 0.4467 0.9845
1.750 0.5254 0.01048 0.00271 -0.0671 0.0061 0.9928
2.000 0.5586 0.01070 0.00303 -0.0686 0.0030 0.9967
2.250 0.5920 0.01108 0.00357 -0.0702 0.0029 0.9999
2.500 0.6108 0.01159 0.00420 -0.0686 0.0031 1.0000
2.750 0.6290 0.01213 0.00492 -0.0667 0.0038 1.0000
3.000 0.6450 0.01282 0.00573 -0.0644 0.0049 1.0000
3.250 0.6603 0.01353 0.00655 -0.0619 0.0089 1.0000
3.500 0.6750 0.01425 0.00735 -0.0592 0.0129 1.0000
3.750 0.6758 0.01587 0.00896 -0.0545 0.0114 1.0000
4.000 0.6929 0.01646 0.00966 -0.0523 0.0098 1.0000
4.250 0.7096 0.01713 0.01038 -0.0501 0.0081 1.0000
4.500 0.7232 0.01841 0.01163 -0.0475 0.0069 1.0000
5.000 0.7735 0.02273 0.01581 -0.0466 0.0033 1.0000
5.250 0.7967 0.02300 0.01627 -0.0447 0.0030 1.0000
5.500 0.8262 0.02433 0.01770 -0.0441 0.0025 1.0000
5.750 0.8549 0.02625 0.01970 -0.0436 0.0021 1.0000
6.000 0.8805 0.02830 0.02186 -0.0429 0.0019 1.0000
6.250 0.9021 0.03145 0.02512 -0.0421 0.0017 1.0000
6.500 0.9220 0.03522 0.02901 -0.0410 0.0016 1.0000
6.750 0.9409 0.03747 0.03145 -0.0390 0.0016 1.0000
7.000 0.9586 0.03967 0.03387 -0.0368 0.0016 1.0000
7.250 0.9762 0.04133 0.03577 -0.0342 0.0015 1.0000
7.500 0.9936 0.04313 0.03780 -0.0317 0.0015 1.0000
7.750 1.0098 0.04516 0.04003 -0.0292 0.0014 1.0000
8.000 1.0235 0.04758 0.04265 -0.0266 0.0013 1.0000
8.250 1.0332 0.05016 0.04544 -0.0240 0.0012 1.0000
8.500 1.0395 0.05297 0.04846 -0.0212 0.0012 1.0000
8.750 1.0432 0.05588 0.05156 -0.0182 0.0011 1.0000
9.000 1.0447 0.05873 0.05460 -0.0153 0.0011 1.0000
9.250 1.0434 0.06152 0.05757 -0.0124 0.0011 1.0000
9.500 1.0383 0.06434 0.06057 -0.0092 0.0010 1.0000
9.750 1.0303 0.06690 0.06328 -0.0061 0.0010 1.0000
10.000 1.0160 0.06933 0.06584 -0.0022 0.0010 1.0000
10.250 1.0006 0.07191 0.06855 0.0008 0.0010 1.0000
10.500 0.9846 0.07474 0.07150 0.0028 0.0010 1.0000
10.750 0.9674 0.07796 0.07484 0.0038 0.0010 1.0000
11.000 0.9483 0.08172 0.07872 0.0040 0.0010 1.0000
11.250 0.9301 0.08578 0.08289 0.0034 0.0010 1.0000
11.500 0.9135 0.09004 0.08725 0.0020 0.0010 1.0000
11.750 0.8942 0.09514 0.09245 0.0000 0.0010 1.0000
12.000 0.8768 0.10048 0.09788 -0.0026 0.0011 1.0000
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Polar data table (+)
Polar graphs
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