GOE 585 AIRFOIL (goe585-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
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Airfoil: GOE 585 AIRFOIL (goe585-il) Reynolds number: 100,000 Max Cl/Cd: 52.66 at α=3.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe585-il-100000.txt Download as CSV file: xf-goe585-il-100000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 585 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 -0.3947 0.09916 0.09410 -0.0255 1.0000 0.0988
-8.250 -0.4172 0.09853 0.09362 -0.0284 1.0000 0.0999
-8.000 -0.4342 0.09686 0.09205 -0.0343 1.0000 0.1004
-7.750 -0.3981 0.08987 0.08501 -0.0253 1.0000 0.1035
-7.500 -0.3941 0.08717 0.08236 -0.0239 1.0000 0.1067
-7.250 -0.3988 0.08483 0.08010 -0.0238 1.0000 0.1102
-7.000 -0.4123 0.08273 0.07808 -0.0297 1.0000 0.1141
-6.750 -0.4177 0.07891 0.07428 -0.0329 1.0000 0.1159
-6.500 -0.4057 0.07600 0.07145 -0.0259 1.0000 0.1194
-6.250 -0.4029 0.07348 0.06895 -0.0251 1.0000 0.1237
-6.000 -0.4078 0.07014 0.06545 -0.0335 1.0000 0.1307
-5.750 -0.4017 0.06683 0.06226 -0.0289 1.0000 0.1325
-5.500 -0.3957 0.06442 0.05989 -0.0263 1.0000 0.1358
-5.250 -0.3878 0.06105 0.05626 -0.0317 1.0000 0.1462
-5.000 -0.3814 0.05820 0.05354 -0.0282 1.0000 0.1486
-4.750 -0.3717 0.05604 0.05135 -0.0272 1.0000 0.1567
-4.500 -0.3624 0.05309 0.04835 -0.0271 1.0000 0.1653
-4.250 -0.3495 0.05043 0.04555 -0.0280 1.0000 0.1784
-4.000 -0.3350 0.04824 0.04319 -0.0285 1.0000 0.1924
-3.750 -0.3243 0.04584 0.04087 -0.0263 1.0000 0.1968
-3.500 -0.2723 0.03530 0.02854 -0.0345 1.0000 0.1145
-3.250 -0.2534 0.03239 0.02559 -0.0340 1.0000 0.1113
-3.000 -0.2256 0.02870 0.02099 -0.0338 1.0000 0.0991
-2.750 -0.2035 0.02672 0.01873 -0.0332 1.0000 0.0976
-2.500 -0.1813 0.02518 0.01685 -0.0324 1.0000 0.0970
-2.250 -0.1595 0.02418 0.01549 -0.0315 1.0000 0.0990
-2.000 -0.1380 0.02321 0.01423 -0.0306 1.0000 0.1012
-1.750 -0.1048 0.02216 0.01309 -0.0320 0.9963 0.1039
-1.500 -0.0539 0.02138 0.01219 -0.0363 0.9864 0.1100
-1.250 -0.0047 0.02059 0.01138 -0.0403 0.9757 0.1206
-1.000 0.0405 0.01992 0.01076 -0.0436 0.9658 0.1442
-0.750 0.0854 0.01870 0.01008 -0.0468 0.9580 0.2237
-0.500 0.1499 0.01608 0.00988 -0.0535 0.9543 1.0000
-0.250 0.1981 0.01622 0.00967 -0.0574 0.9441 1.0000
0.000 0.2405 0.01627 0.00950 -0.0601 0.9321 1.0000
0.250 0.2842 0.01627 0.00934 -0.0629 0.9204 1.0000
0.500 0.3337 0.01613 0.00908 -0.0667 0.9100 1.0000
0.750 0.3868 0.01582 0.00868 -0.0709 0.8998 1.0000
1.000 0.4299 0.01554 0.00835 -0.0731 0.8865 1.0000
1.250 0.4716 0.01521 0.00798 -0.0749 0.8724 1.0000
1.500 0.5111 0.01485 0.00758 -0.0760 0.8570 1.0000
1.750 0.5484 0.01446 0.00718 -0.0767 0.8400 1.0000
2.000 0.5794 0.01418 0.00687 -0.0761 0.8194 1.0000
2.250 0.6095 0.01390 0.00655 -0.0753 0.7961 1.0000
2.500 0.6374 0.01370 0.00631 -0.0741 0.7687 1.0000
2.750 0.6649 0.01358 0.00609 -0.0728 0.7373 1.0000
3.000 0.6906 0.01359 0.00595 -0.0713 0.7003 1.0000
3.250 0.7142 0.01374 0.00592 -0.0695 0.6577 1.0000
3.500 0.7359 0.01401 0.00598 -0.0676 0.6109 1.0000
3.750 0.7572 0.01438 0.00610 -0.0658 0.5659 1.0000
4.000 0.7787 0.01485 0.00632 -0.0641 0.5265 1.0000
4.250 0.8008 0.01539 0.00664 -0.0627 0.4944 1.0000
4.500 0.8234 0.01599 0.00705 -0.0615 0.4669 1.0000
4.750 0.8467 0.01664 0.00753 -0.0606 0.4443 1.0000
5.000 0.8705 0.01730 0.00810 -0.0597 0.4245 1.0000
5.250 0.8945 0.01798 0.00872 -0.0589 0.4072 1.0000
5.500 0.9187 0.01868 0.00936 -0.0582 0.3916 1.0000
5.750 0.9430 0.01939 0.01007 -0.0575 0.3777 1.0000
6.000 0.9671 0.02011 0.01082 -0.0568 0.3648 1.0000
6.250 0.9909 0.02082 0.01156 -0.0560 0.3519 1.0000
6.500 1.0135 0.02143 0.01218 -0.0550 0.3377 1.0000
6.750 1.0356 0.02200 0.01277 -0.0540 0.3233 1.0000
7.000 1.0581 0.02260 0.01340 -0.0530 0.3103 1.0000
7.250 1.0805 0.02319 0.01403 -0.0521 0.2978 1.0000
7.500 1.1013 0.02376 0.01475 -0.0508 0.2847 1.0000
7.750 1.1217 0.02431 0.01543 -0.0495 0.2712 1.0000
8.000 1.1412 0.02480 0.01601 -0.0481 0.2564 1.0000
8.250 1.1594 0.02518 0.01646 -0.0464 0.2399 1.0000
8.500 1.1736 0.02543 0.01681 -0.0441 0.2199 1.0000
8.750 1.1854 0.02561 0.01705 -0.0416 0.1971 1.0000
9.000 1.1956 0.02610 0.01761 -0.0388 0.1713 1.0000
9.250 1.2041 0.02692 0.01840 -0.0360 0.1434 1.0000
9.500 1.2115 0.02816 0.01958 -0.0331 0.1176 1.0000
9.750 1.2167 0.02986 0.02109 -0.0301 0.0980 1.0000
10.000 1.2233 0.03180 0.02299 -0.0273 0.0842 1.0000
10.250 1.2331 0.03413 0.02514 -0.0252 0.0752 1.0000
10.500 1.2430 0.03594 0.02720 -0.0228 0.0685 1.0000
10.750 1.2584 0.03858 0.02974 -0.0217 0.0628 1.0000
11.000 1.2701 0.04078 0.03224 -0.0197 0.0595 1.0000
11.250 1.2822 0.04319 0.03489 -0.0180 0.0567 1.0000
11.500 1.2980 0.04599 0.03770 -0.0172 0.0539 1.0000
11.750 1.3046 0.04961 0.04156 -0.0155 0.0519 1.0000
12.000 1.2978 0.05188 0.04418 -0.0120 0.0509 1.0000
12.250 1.2897 0.05466 0.04730 -0.0091 0.0501 1.0000
12.500 1.2783 0.05783 0.05078 -0.0065 0.0495 1.0000
12.750 1.2649 0.06139 0.05464 -0.0047 0.0494 1.0000
13.000 1.2474 0.06543 0.05896 -0.0036 0.0493 1.0000
13.250 1.2274 0.07001 0.06382 -0.0035 0.0493 1.0000
13.500 1.2049 0.07526 0.06932 -0.0045 0.0495 1.0000
13.750 1.1807 0.08129 0.07558 -0.0067 0.0499 1.0000
14.000 1.1538 0.08836 0.08287 -0.0103 0.0503 1.0000
14.250 1.1262 0.09647 0.09115 -0.0152 0.0509 1.0000
14.500 1.0967 0.10585 0.10068 -0.0212 0.0515 1.0000
14.750 1.0694 0.11596 0.11087 -0.0274 0.0523 1.0000
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Polar data table (+)
Polar graphs
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