GOE 499 AIRFOIL (goe499-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
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Airfoil: GOE 499 AIRFOIL (goe499-il) Reynolds number: 100,000 Max Cl/Cd: 71.22 at α=5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe499-il-100000.txt Download as CSV file: xf-goe499-il-100000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 499 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
------ -------- --------- --------- -------- -------- --------
-7.000 -0.3794 0.10651 0.10219 -0.0170 1.0000 0.0757
-6.750 -0.3922 0.10573 0.10148 -0.0150 1.0000 0.0775
-6.500 -0.4049 0.10535 0.10117 -0.0151 1.0000 0.0790
-6.250 -0.4102 0.10511 0.10100 -0.0215 1.0000 0.0800
-6.000 -0.4092 0.10090 0.09683 -0.0208 1.0000 0.0809
-5.750 -0.4061 0.09703 0.09297 -0.0144 1.0000 0.0830
-5.500 -0.4030 0.09451 0.09047 -0.0131 1.0000 0.0858
-5.250 -0.3987 0.09207 0.08804 -0.0144 1.0000 0.0894
-5.000 -0.3703 0.08870 0.08463 -0.0327 1.0000 0.0943
-4.750 -0.3768 0.08539 0.08137 -0.0241 1.0000 0.0959
-4.500 -0.3730 0.08301 0.07901 -0.0208 1.0000 0.1002
-4.250 -0.3326 0.07836 0.07425 -0.0374 1.0000 0.1090
-4.000 -0.3327 0.07566 0.07157 -0.0321 1.0000 0.1111
-3.750 -0.2779 0.07089 0.06660 -0.0497 1.0000 0.1230
-3.500 -0.2780 0.06819 0.06396 -0.0445 1.0000 0.1247
-3.250 -0.2660 0.06584 0.06161 -0.0433 1.0000 0.1301
-3.000 -0.2115 0.06164 0.05726 -0.0543 0.9946 0.1443
-2.750 -0.1450 0.05707 0.05247 -0.0675 0.9855 0.1588
-2.500 -0.0849 0.05314 0.04835 -0.0785 0.9774 0.1743
-2.250 -0.0225 0.04959 0.04459 -0.0894 0.9713 0.1977
-2.000 0.0377 0.04677 0.04146 -0.0999 0.9646 0.2238
-1.750 0.1933 0.03028 0.02207 -0.1301 0.9699 0.0934
-1.500 0.2294 0.02887 0.02032 -0.1322 0.9634 0.0949
-1.250 0.2718 0.02817 0.01952 -0.1354 0.9583 0.1006
-1.000 0.3085 0.02747 0.01858 -0.1372 0.9520 0.1032
-0.750 0.3476 0.02696 0.01785 -0.1393 0.9459 0.1076
-0.500 0.3867 0.02651 0.01751 -0.1416 0.9402 0.1178
-0.250 0.4221 0.02611 0.01725 -0.1431 0.9328 0.1456
0.000 0.4671 0.02590 0.01743 -0.1466 0.9275 0.2877
0.250 0.4983 0.02589 0.01760 -0.1475 0.9183 0.3582
0.500 0.5358 0.02585 0.01779 -0.1494 0.9112 0.4323
0.750 0.5723 0.02551 0.01797 -0.1510 0.9036 0.5871
1.000 0.6003 0.02482 0.01778 -0.1507 0.8943 1.0000
1.250 0.6448 0.02489 0.01764 -0.1536 0.8877 1.0000
1.500 0.6755 0.02506 0.01770 -0.1540 0.8766 1.0000
1.750 0.7113 0.02515 0.01773 -0.1553 0.8675 1.0000
2.000 0.7535 0.02504 0.01758 -0.1576 0.8602 1.0000
2.250 0.7852 0.02508 0.01760 -0.1581 0.8489 1.0000
2.500 0.8341 0.02465 0.01722 -0.1612 0.8438 1.0000
2.750 0.8666 0.02452 0.01714 -0.1616 0.8322 1.0000
3.000 0.8992 0.02439 0.01707 -0.1620 0.8208 1.0000
3.250 0.9503 0.02362 0.01644 -0.1652 0.8167 1.0000
3.500 0.9823 0.02333 0.01626 -0.1651 0.8043 1.0000
3.750 1.0163 0.02288 0.01594 -0.1653 0.7922 1.0000
4.000 1.0534 0.02211 0.01535 -0.1656 0.7801 1.0000
4.250 1.1066 0.01907 0.01243 -0.1658 0.7600 1.0000
4.500 1.1328 0.01788 0.01132 -0.1631 0.7277 1.0000
4.750 1.1573 0.01690 0.01032 -0.1601 0.6779 1.0000
5.000 1.1794 0.01656 0.00948 -0.1564 0.5720 1.0000
5.250 1.1813 0.01824 0.00994 -0.1505 0.4100 1.0000
5.500 1.1661 0.02228 0.01179 -0.1442 0.1009 1.0000
5.750 1.1787 0.02396 0.01327 -0.1414 0.0679 1.0000
6.250 1.2042 0.02667 0.01619 -0.1359 0.0570 1.0000
6.500 1.2139 0.02803 0.01766 -0.1328 0.0532 1.0000
6.750 1.2197 0.02969 0.01934 -0.1293 0.0500 1.0000
7.000 1.2299 0.03184 0.02148 -0.1265 0.0482 1.0000
7.250 1.2578 0.03386 0.02352 -0.1260 0.0475 1.0000
7.500 1.3029 0.03665 0.02634 -0.1280 0.0472 1.0000
7.750 1.3506 0.04028 0.03006 -0.1302 0.0477 1.0000
8.000 1.3809 0.04313 0.03321 -0.1298 0.0472 1.0000
8.250 1.4084 0.04656 0.03702 -0.1292 0.0473 1.0000
8.500 1.4377 0.04952 0.04031 -0.1282 0.0502 1.0000
12.250 0.9600 0.15716 0.15369 -0.1205 0.1693 1.0000
12.500 0.9869 0.16047 0.15705 -0.1177 0.1659 1.0000
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Polar data table (+)
Polar graphs
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