GOE 617 AIRFOIL (goe617-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: GOE 617 AIRFOIL (goe617-il) Reynolds number: 100,000 Max Cl/Cd: 49.43 at α=6.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe617-il-100000.txt Download as CSV file: xf-goe617-il-100000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 617 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
------ -------- --------- --------- -------- -------- --------
-9.750 -0.4089 0.09595 0.09071 -0.0353 1.0000 0.1778
-9.500 -0.4779 0.09178 0.08678 -0.0414 1.0000 0.1859
-9.250 -0.4523 0.08803 0.08302 -0.0382 1.0000 0.1886
-9.000 -0.4324 0.08582 0.08080 -0.0355 1.0000 0.1927
-8.750 -0.4470 0.08270 0.07777 -0.0354 1.0000 0.1982
-8.500 -0.5400 0.07999 0.07529 -0.0333 1.0000 0.2019
-8.250 -0.7113 0.05824 0.05256 -0.0264 1.0000 0.1090
-8.000 -0.7138 0.05505 0.04932 -0.0233 1.0000 0.1073
-7.750 -0.7226 0.05162 0.04570 -0.0193 1.0000 0.1048
-7.500 -0.7599 0.04530 0.03825 -0.0110 1.0000 0.0973
-7.250 -0.7581 0.04274 0.03537 -0.0075 1.0000 0.0972
-7.000 -0.7488 0.04046 0.03308 -0.0054 1.0000 0.0989
-6.750 -0.7383 0.03899 0.03155 -0.0031 1.0000 0.1010
-6.500 -0.7303 0.03703 0.02931 -0.0003 1.0000 0.1019
-6.250 -0.7206 0.03514 0.02711 0.0025 1.0000 0.1031
-6.000 -0.7090 0.03352 0.02518 0.0049 1.0000 0.1047
-5.750 -0.6968 0.03221 0.02353 0.0074 1.0000 0.1076
-5.500 -0.6762 0.03104 0.02188 0.0085 0.9979 0.1102
-5.250 -0.6364 0.02917 0.02000 0.0055 0.9914 0.1143
-5.000 -0.5961 0.02829 0.01894 0.0027 0.9840 0.1209
-4.750 -0.5578 0.02713 0.01766 0.0005 0.9762 0.1276
-4.500 -0.5132 0.02647 0.01690 -0.0030 0.9694 0.1369
-4.250 -0.4791 0.02566 0.01618 -0.0045 0.9600 0.1468
-4.000 -0.4389 0.02504 0.01556 -0.0071 0.9525 0.1589
-3.750 -0.4013 0.02447 0.01502 -0.0092 0.9435 0.1733
-3.500 -0.3665 0.02391 0.01456 -0.0108 0.9346 0.1906
-3.250 -0.3273 0.02318 0.01399 -0.0131 0.9266 0.2174
-3.000 -0.3020 0.02217 0.01341 -0.0129 0.9165 0.2613
-2.750 -0.2775 0.02031 0.01310 -0.0122 0.9092 0.5003
-2.500 -0.2000 0.01987 0.01393 -0.0191 0.9084 0.8082
-2.250 -0.1362 0.02063 0.01460 -0.0243 0.9012 0.8775
-2.000 -0.0473 0.02147 0.01522 -0.0340 0.8984 0.9126
-1.750 0.0466 0.02201 0.01554 -0.0450 0.8968 0.9389
-1.500 0.1638 0.02196 0.01529 -0.0608 0.8977 0.9607
-1.250 0.2589 0.02126 0.01445 -0.0734 0.8952 0.9769
-1.000 0.3467 0.02021 0.01328 -0.0850 0.8917 0.9895
-0.750 0.4088 0.01936 0.01235 -0.0922 0.8801 1.0000
-0.500 0.4393 0.01895 0.01187 -0.0931 0.8659 1.0000
-0.250 0.4672 0.01862 0.01147 -0.0933 0.8516 1.0000
0.000 0.4934 0.01835 0.01112 -0.0932 0.8376 1.0000
0.250 0.5193 0.01809 0.01077 -0.0928 0.8241 1.0000
0.500 0.5380 0.01801 0.01064 -0.0911 0.8087 1.0000
0.750 0.5567 0.01797 0.01054 -0.0895 0.7937 1.0000
1.000 0.5756 0.01794 0.01047 -0.0878 0.7791 1.0000
1.250 0.5950 0.01793 0.01041 -0.0862 0.7651 1.0000
1.500 0.6157 0.01792 0.01035 -0.0848 0.7521 1.0000
1.750 0.6385 0.01785 0.01021 -0.0836 0.7399 1.0000
2.000 0.6588 0.01784 0.01015 -0.0820 0.7260 1.0000
2.250 0.6776 0.01789 0.01016 -0.0802 0.7113 1.0000
2.500 0.6970 0.01792 0.01017 -0.0784 0.6967 1.0000
2.750 0.7168 0.01798 0.01018 -0.0767 0.6824 1.0000
3.000 0.7371 0.01803 0.01018 -0.0750 0.6683 1.0000
3.250 0.7584 0.01811 0.01020 -0.0736 0.6551 1.0000
3.500 0.7817 0.01817 0.01019 -0.0724 0.6425 1.0000
3.750 0.8000 0.01833 0.01034 -0.0705 0.6279 1.0000
4.000 0.8183 0.01850 0.01050 -0.0685 0.6133 1.0000
4.250 0.8370 0.01868 0.01067 -0.0666 0.5989 1.0000
4.500 0.8563 0.01888 0.01086 -0.0649 0.5855 1.0000
4.750 0.8767 0.01906 0.01102 -0.0633 0.5725 1.0000
5.000 0.8966 0.01920 0.01110 -0.0615 0.5581 1.0000
5.250 0.9155 0.01932 0.01119 -0.0595 0.5429 1.0000
5.500 0.9325 0.01946 0.01132 -0.0573 0.5271 1.0000
5.750 0.9486 0.01961 0.01147 -0.0548 0.5110 1.0000
6.000 0.9644 0.01978 0.01165 -0.0524 0.4948 1.0000
6.250 0.9797 0.01996 0.01184 -0.0498 0.4784 1.0000
6.500 0.9944 0.02013 0.01199 -0.0472 0.4611 1.0000
6.750 1.0055 0.02034 0.01224 -0.0439 0.4417 1.0000
7.000 1.0157 0.02056 0.01249 -0.0404 0.4205 1.0000
7.250 1.0255 0.02081 0.01266 -0.0369 0.3976 1.0000
7.500 1.0313 0.02119 0.01299 -0.0327 0.3711 1.0000
7.750 1.0360 0.02169 0.01335 -0.0284 0.3432 1.0000
8.000 1.0398 0.02233 0.01378 -0.0241 0.3165 1.0000
8.250 1.0425 0.02305 0.01442 -0.0197 0.2905 1.0000
8.500 1.0468 0.02384 0.01501 -0.0157 0.2697 1.0000
8.750 1.0516 0.02466 0.01571 -0.0119 0.2510 1.0000
9.000 1.0577 0.02552 0.01646 -0.0083 0.2352 1.0000
9.250 1.0662 0.02643 0.01722 -0.0053 0.2221 1.0000
9.500 1.0764 0.02737 0.01811 -0.0025 0.2104 1.0000
9.750 1.0894 0.02840 0.01912 -0.0003 0.1999 1.0000
10.000 1.1100 0.02962 0.02023 0.0005 0.1905 1.0000
10.250 1.1275 0.03073 0.02134 0.0018 0.1823 1.0000
10.500 1.1550 0.03224 0.02285 0.0013 0.1752 1.0000
10.750 1.1720 0.03347 0.02419 0.0027 0.1693 1.0000
11.000 1.2088 0.03532 0.02589 0.0006 0.1626 1.0000
11.250 1.2149 0.03650 0.02735 0.0038 0.1589 1.0000
11.500 1.2271 0.03789 0.02893 0.0058 0.1548 1.0000
11.750 1.2536 0.03956 0.03057 0.0054 0.1501 1.0000
12.000 1.2665 0.04152 0.03270 0.0070 0.1466 1.0000
12.250 1.2633 0.04291 0.03438 0.0111 0.1437 1.0000
12.500 1.2643 0.04422 0.03585 0.0146 0.1400 1.0000
12.750 1.2935 0.04595 0.03740 0.0133 0.1333 1.0000
13.000 1.2743 0.04708 0.03884 0.0195 0.1314 1.0000
13.250 1.2593 0.04841 0.04041 0.0244 0.1285 1.0000
13.500 1.2590 0.04945 0.04152 0.0271 0.1240 1.0000
13.750 1.2763 0.05088 0.04283 0.0276 0.1189 1.0000
14.000 1.2552 0.05300 0.04528 0.0318 0.1172 1.0000
14.250 1.2366 0.05545 0.04800 0.0349 0.1151 1.0000
14.500 1.2274 0.05766 0.05038 0.0369 0.1121 1.0000
14.750 1.2558 0.05821 0.05070 0.0368 0.1061 1.0000
15.000 1.2273 0.06167 0.05449 0.0392 0.1051 1.0000
15.250 1.1969 0.06585 0.05897 0.0406 0.1041 1.0000
15.500 1.1642 0.07087 0.06426 0.0409 0.1032 1.0000
15.750 1.1268 0.07720 0.07085 0.0400 0.1030 1.0000
16.000 1.0820 0.08548 0.07938 0.0375 0.1040 1.0000
16.250 1.0300 0.09601 0.09010 0.0331 0.1053 1.0000
16.500 1.1034 0.08779 0.08164 0.0381 0.0937 1.0000
16.750 1.0594 0.09697 0.09101 0.0343 0.0940 1.0000
17.000 0.7146 0.18543 0.17938 -0.0164 0.1613 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 617 AIRFOIL (goe617-il)