GOE 502 AIRFOIL (goe502-il) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: GOE 502 AIRFOIL (goe502-il) Reynolds number: 200,000 Max Cl/Cd: 75.95 at α=6.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe502-il-200000.txt Download as CSV file: xf-goe502-il-200000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 502 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.200 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.500 -0.0506 0.10185 0.09795 -0.0916 0.9558 0.0995
-9.250 -0.2968 0.04707 0.04251 -0.1442 0.9239 0.0701
-9.000 -0.2991 0.04181 0.03694 -0.1481 0.9119 0.0696
-8.750 -0.2784 0.03639 0.03103 -0.1547 0.9076 0.0693
-8.500 -0.2473 0.03252 0.02665 -0.1600 0.9053 0.0694
-8.250 -0.2397 0.03085 0.02468 -0.1582 0.8938 0.0697
-8.000 -0.2077 0.02866 0.02205 -0.1609 0.8904 0.0705
-7.750 -0.1716 0.02659 0.01987 -0.1635 0.8881 0.0714
-7.500 -0.1523 0.02557 0.01877 -0.1624 0.8802 0.0721
-7.250 -0.1242 0.02444 0.01752 -0.1629 0.8746 0.0730
-7.000 -0.0889 0.02323 0.01617 -0.1646 0.8712 0.0741
-6.750 -0.0513 0.02209 0.01489 -0.1666 0.8685 0.0754
-6.500 -0.0358 0.02158 0.01428 -0.1644 0.8592 0.0765
-6.250 -0.0049 0.02073 0.01329 -0.1651 0.8543 0.0781
-6.000 0.0302 0.01977 0.01239 -0.1665 0.8509 0.0803
-5.750 0.0538 0.01928 0.01188 -0.1657 0.8440 0.0821
-5.500 0.0795 0.01872 0.01128 -0.1652 0.8373 0.0841
-5.250 0.1133 0.01802 0.01051 -0.1662 0.8330 0.0865
-5.000 0.1420 0.01751 0.01003 -0.1664 0.8277 0.0898
-4.750 0.1645 0.01719 0.00968 -0.1652 0.8198 0.0937
-4.500 0.1973 0.01660 0.00911 -0.1660 0.8149 0.0990
-4.250 0.2263 0.01618 0.00866 -0.1661 0.8090 0.1061
-4.000 0.2503 0.01581 0.00834 -0.1653 0.8013 0.1173
-3.750 0.2830 0.01526 0.00789 -0.1661 0.7961 0.1464
-3.500 0.3102 0.01510 0.00779 -0.1659 0.7892 0.1785
-3.250 0.3380 0.01506 0.00770 -0.1656 0.7817 0.1986
-3.000 0.3724 0.01497 0.00753 -0.1665 0.7761 0.2144
-2.750 0.3961 0.01504 0.00760 -0.1656 0.7678 0.2264
-2.500 0.4265 0.01502 0.00745 -0.1657 0.7607 0.2386
-2.250 0.4565 0.01499 0.00742 -0.1659 0.7537 0.2489
-2.000 0.4818 0.01498 0.00736 -0.1652 0.7450 0.2591
-1.750 0.5144 0.01499 0.00730 -0.1658 0.7385 0.2713
-1.500 0.5376 0.01502 0.00735 -0.1647 0.7295 0.2816
-1.250 0.5670 0.01503 0.00728 -0.1647 0.7218 0.2931
-1.000 0.5941 0.01504 0.00729 -0.1644 0.7140 0.3041
-0.750 0.6197 0.01513 0.00730 -0.1637 0.7054 0.3165
-0.500 0.6514 0.01512 0.00728 -0.1642 0.6987 0.3299
-0.250 0.6722 0.01517 0.00737 -0.1627 0.6894 0.3419
0.000 0.7007 0.01521 0.00736 -0.1626 0.6820 0.3553
0.250 0.7262 0.01524 0.00742 -0.1620 0.6745 0.3667
0.500 0.7510 0.01529 0.00743 -0.1612 0.6662 0.3788
0.750 0.7811 0.01527 0.00739 -0.1614 0.6599 0.3910
1.000 0.8023 0.01534 0.00751 -0.1600 0.6515 0.4032
1.250 0.8289 0.01535 0.00752 -0.1595 0.6445 0.4169
1.500 0.8566 0.01539 0.00756 -0.1593 0.6382 0.4309
1.750 0.8786 0.01544 0.00766 -0.1581 0.6302 0.4453
2.000 0.9062 0.01546 0.00765 -0.1578 0.6240 0.4623
2.250 0.9312 0.01551 0.00776 -0.1572 0.6176 0.4796
2.500 0.9543 0.01553 0.00786 -0.1561 0.6103 0.4989
2.750 0.9820 0.01549 0.00785 -0.1560 0.6045 0.5225
3.000 1.0049 0.01552 0.00803 -0.1549 0.5981 0.5500
3.250 1.0281 0.01549 0.00818 -0.1539 0.5915 0.5903
3.500 1.0530 0.01534 0.00827 -0.1531 0.5861 0.6758
3.750 1.0868 0.01503 0.00842 -0.1540 0.5796 1.0000
4.000 1.1111 0.01525 0.00857 -0.1533 0.5731 1.0000
4.250 1.1400 0.01547 0.00866 -0.1535 0.5678 1.0000
4.500 1.1646 0.01577 0.00892 -0.1528 0.5620 1.0000
4.750 1.1870 0.01602 0.00915 -0.1518 0.5554 1.0000
5.000 1.2154 0.01622 0.00923 -0.1517 0.5490 1.0000
5.250 1.2367 0.01652 0.00953 -0.1505 0.5422 1.0000
5.500 1.2591 0.01676 0.00975 -0.1494 0.5353 1.0000
5.750 1.2895 0.01703 0.00988 -0.1498 0.5292 1.0000
6.000 1.3058 0.01732 0.01025 -0.1476 0.5222 1.0000
6.250 1.3286 0.01758 0.01048 -0.1467 0.5156 1.0000
6.500 1.3603 0.01791 0.01069 -0.1474 0.5099 1.0000
6.750 1.3728 0.01821 0.01110 -0.1445 0.5030 1.0000
7.000 1.3946 0.01846 0.01132 -0.1434 0.4963 1.0000
7.250 1.4222 0.01879 0.01158 -0.1433 0.4902 1.0000
7.500 1.4342 0.01909 0.01199 -0.1404 0.4834 1.0000
7.750 1.4554 0.01936 0.01224 -0.1392 0.4772 1.0000
8.000 1.4813 0.01970 0.01253 -0.1389 0.4714 1.0000
8.250 1.4910 0.02004 0.01298 -0.1356 0.4648 1.0000
8.500 1.5090 0.02031 0.01326 -0.1338 0.4588 1.0000
8.750 1.5338 0.02064 0.01354 -0.1333 0.4530 1.0000
9.000 1.5374 0.02101 0.01404 -0.1290 0.4466 1.0000
9.250 1.5513 0.02128 0.01431 -0.1265 0.4398 1.0000
9.500 1.5647 0.02163 0.01467 -0.1241 0.4327 1.0000
9.750 1.5687 0.02203 0.01515 -0.1201 0.4248 1.0000
10.000 1.5861 0.02233 0.01537 -0.1183 0.4173 1.0000
10.250 1.5854 0.02293 0.01612 -0.1139 0.4092 1.0000
10.500 1.5974 0.02333 0.01646 -0.1115 0.4010 1.0000
10.750 1.5979 0.02407 0.01733 -0.1075 0.3921 1.0000
11.000 1.6069 0.02464 0.01785 -0.1049 0.3832 1.0000
11.250 1.6065 0.02558 0.01892 -0.1012 0.3732 1.0000
11.500 1.6117 0.02642 0.01973 -0.0984 0.3632 1.0000
11.750 1.6122 0.02753 0.02090 -0.0952 0.3519 1.0000
12.000 1.6132 0.02875 0.02215 -0.0922 0.3402 1.0000
12.250 1.6138 0.03009 0.02347 -0.0894 0.3282 1.0000
12.500 1.6118 0.03169 0.02506 -0.0865 0.3150 1.0000
12.750 1.6082 0.03354 0.02694 -0.0838 0.3004 1.0000
13.000 1.6036 0.03562 0.02900 -0.0813 0.2854 1.0000
13.250 1.5979 0.03796 0.03131 -0.0789 0.2700 1.0000
13.500 1.5918 0.04048 0.03379 -0.0768 0.2552 1.0000
13.750 1.5858 0.04315 0.03642 -0.0749 0.2417 1.0000
14.000 1.5804 0.04589 0.03910 -0.0733 0.2301 1.0000
14.250 1.5758 0.04866 0.04183 -0.0718 0.2194 1.0000
14.500 1.5738 0.05130 0.04449 -0.0707 0.2100 1.0000
15.000 1.5696 0.05680 0.04997 -0.0687 0.1940 1.0000
15.250 1.5677 0.05955 0.05265 -0.0678 0.1874 1.0000
15.500 1.5676 0.06232 0.05551 -0.0672 0.1803 1.0000
15.750 1.5667 0.06510 0.05824 -0.0665 0.1739 1.0000
16.000 1.5666 0.06794 0.06116 -0.0661 0.1675 1.0000
16.250 1.5655 0.07090 0.06412 -0.0658 0.1612 1.0000
16.500 1.5653 0.07379 0.06703 -0.0654 0.1549 1.0000
16.750 1.5627 0.07711 0.07042 -0.0654 0.1485 1.0000
17.000 1.5620 0.08012 0.07340 -0.0653 0.1423 1.0000
17.250 1.5580 0.08380 0.07718 -0.0655 0.1358 1.0000
17.500 1.5565 0.08697 0.08031 -0.0656 0.1297 1.0000
17.750 1.5524 0.09078 0.08424 -0.0661 0.1236 1.0000
18.000 1.5512 0.09401 0.08740 -0.0664 0.1181 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 502 AIRFOIL (goe502-il)