GOE 481A AIRFOIL (goe481a-il) Xfoil prediction polar at RE=200,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: GOE 481A AIRFOIL (goe481a-il) Reynolds number: 200,000 Max Cl/Cd: 59.03 at α=6° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe481a-il-200000-n5.txt Download as CSV file: xf-goe481a-il-200000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 481A 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 = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.750 -0.1182 0.09337 0.08904 -0.0858 0.9463 0.0543
-8.500 -0.1259 0.08944 0.08513 -0.0866 0.9404 0.0547
-8.250 -0.1247 0.08558 0.08127 -0.0886 0.9358 0.0546
-7.750 -0.3469 0.04212 0.03696 -0.1084 0.9040 0.0562
-7.500 -0.3558 0.03685 0.03121 -0.1070 0.8970 0.0567
-7.250 -0.3486 0.03434 0.02843 -0.1053 0.8904 0.0572
-7.000 -0.3289 0.03199 0.02580 -0.1053 0.8871 0.0578
-6.750 -0.3049 0.02986 0.02337 -0.1057 0.8849 0.0584
-6.500 -0.3016 0.02875 0.02206 -0.1016 0.8748 0.0588
-6.250 -0.2783 0.02725 0.02029 -0.1012 0.8710 0.0596
-6.000 -0.2514 0.02576 0.01851 -0.1014 0.8684 0.0606
-5.750 -0.2400 0.02495 0.01747 -0.0983 0.8593 0.0612
-5.500 -0.2157 0.02396 0.01633 -0.0976 0.8542 0.0619
-5.250 -0.1862 0.02307 0.01539 -0.0977 0.8511 0.0624
-5.000 -0.1694 0.02258 0.01488 -0.0955 0.8423 0.0630
-4.750 -0.1448 0.02190 0.01414 -0.0946 0.8358 0.0636
-4.500 -0.1128 0.02108 0.01324 -0.0951 0.8319 0.0644
-4.250 -0.0975 0.02067 0.01276 -0.0925 0.8201 0.0651
-4.000 -0.0657 0.01991 0.01190 -0.0929 0.8146 0.0660
-3.750 -0.0444 0.01944 0.01133 -0.0913 0.8032 0.0670
-3.500 -0.0099 0.01870 0.01047 -0.0922 0.7960 0.0683
-3.250 0.0164 0.01821 0.00999 -0.0916 0.7839 0.0693
-3.000 0.0544 0.01756 0.00930 -0.0932 0.7740 0.0708
-2.750 0.0914 0.01697 0.00862 -0.0946 0.7605 0.0724
-2.500 0.1315 0.01638 0.00790 -0.0966 0.7447 0.0740
-2.250 0.1724 0.01585 0.00728 -0.0988 0.7270 0.0757
-2.000 0.2112 0.01548 0.00681 -0.1008 0.7073 0.0779
-1.750 0.2449 0.01524 0.00643 -0.1016 0.6881 0.0807
-1.500 0.2740 0.01509 0.00618 -0.1016 0.6702 0.0835
-1.250 0.3007 0.01501 0.00602 -0.1012 0.6542 0.0869
-1.000 0.3265 0.01495 0.00586 -0.1005 0.6399 0.0906
-0.750 0.3492 0.01495 0.00585 -0.0993 0.6266 0.0948
-0.500 0.3727 0.01497 0.00582 -0.0981 0.6147 0.1005
-0.250 0.3955 0.01504 0.00583 -0.0969 0.6027 0.1083
0.000 0.4180 0.01514 0.00589 -0.0956 0.5921 0.1154
0.250 0.4410 0.01525 0.00595 -0.0944 0.5824 0.1226
0.500 0.4635 0.01535 0.00597 -0.0931 0.5733 0.1299
0.750 0.4854 0.01545 0.00603 -0.0917 0.5643 0.1351
1.000 0.5079 0.01558 0.00606 -0.0904 0.5556 0.1415
1.250 0.5288 0.01562 0.00613 -0.0889 0.5470 0.1462
1.500 0.5510 0.01574 0.00616 -0.0876 0.5390 0.1516
1.750 0.5719 0.01579 0.00620 -0.0861 0.5314 0.1559
2.000 0.5929 0.01583 0.00626 -0.0846 0.5239 0.1606
2.250 0.6148 0.01595 0.00632 -0.0833 0.5170 0.1666
2.500 0.6345 0.01598 0.00640 -0.0815 0.5091 0.1730
2.750 0.6538 0.01607 0.00647 -0.0797 0.5005 0.1792
3.000 0.6729 0.01618 0.00655 -0.0779 0.4917 0.1841
3.250 0.6913 0.01624 0.00662 -0.0759 0.4830 0.1879
3.500 0.7104 0.01634 0.00671 -0.0741 0.4751 0.1917
3.750 0.7296 0.01643 0.00682 -0.0723 0.4674 0.1960
4.000 0.7492 0.01657 0.00693 -0.0706 0.4607 0.2009
4.250 0.7684 0.01668 0.00707 -0.0688 0.4535 0.2070
4.500 0.7865 0.01681 0.00722 -0.0669 0.4449 0.2148
4.750 0.8052 0.01696 0.00738 -0.0651 0.4381 0.2284
5.000 0.8246 0.01703 0.00759 -0.0635 0.4306 0.2639
5.500 1.0031 0.01701 0.00901 -0.0909 0.4056 1.0000
5.750 1.0204 0.01732 0.00924 -0.0889 0.3979 1.0000
6.000 1.0390 0.01760 0.00950 -0.0872 0.3903 1.0000
6.250 1.0562 0.01793 0.00975 -0.0852 0.3835 1.0000
6.500 1.0750 0.01823 0.01004 -0.0836 0.3773 1.0000
6.750 1.0932 0.01855 0.01034 -0.0818 0.3707 1.0000
7.000 1.1102 0.01893 0.01065 -0.0799 0.3647 1.0000
7.250 1.1288 0.01926 0.01098 -0.0784 0.3588 1.0000
7.500 1.1466 0.01962 0.01134 -0.0767 0.3526 1.0000
7.750 1.1628 0.02004 0.01172 -0.0747 0.3466 1.0000
8.000 1.1804 0.02043 0.01211 -0.0731 0.3410 1.0000
8.250 1.1976 0.02084 0.01253 -0.0713 0.3347 1.0000
8.500 1.2124 0.02133 0.01298 -0.0693 0.3283 1.0000
8.750 1.2283 0.02179 0.01345 -0.0675 0.3217 1.0000
9.000 1.2426 0.02231 0.01396 -0.0654 0.3136 1.0000
9.250 1.2552 0.02291 0.01453 -0.0632 0.3064 1.0000
9.500 1.2699 0.02344 0.01509 -0.0613 0.2983 1.0000
9.750 1.2803 0.02416 0.01575 -0.0589 0.2900 1.0000
10.000 1.2953 0.02473 0.01637 -0.0572 0.2828 1.0000
10.250 1.3061 0.02548 0.01711 -0.0550 0.2743 1.0000
10.500 1.3186 0.02619 0.01784 -0.0530 0.2667 1.0000
10.750 1.3289 0.02702 0.01867 -0.0509 0.2583 1.0000
11.000 1.3395 0.02787 0.01953 -0.0489 0.2502 1.0000
11.250 1.3486 0.02882 0.02048 -0.0468 0.2418 1.0000
11.500 1.3572 0.02984 0.02151 -0.0447 0.2333 1.0000
11.750 1.3641 0.03100 0.02266 -0.0426 0.2250 1.0000
12.000 1.3710 0.03221 0.02387 -0.0405 0.2163 1.0000
12.250 1.3764 0.03356 0.02522 -0.0385 0.2094 1.0000
12.500 1.3831 0.03489 0.02657 -0.0367 0.2032 1.0000
12.750 1.3881 0.03637 0.02806 -0.0349 0.1972 1.0000
13.000 1.3912 0.03806 0.02975 -0.0330 0.1923 1.0000
13.250 1.3977 0.03955 0.03129 -0.0316 0.1874 1.0000
13.500 1.4023 0.04123 0.03300 -0.0301 0.1832 1.0000
13.750 1.4052 0.04311 0.03491 -0.0287 0.1796 1.0000
14.000 1.4088 0.04498 0.03682 -0.0274 0.1763 1.0000
14.250 1.4149 0.04670 0.03861 -0.0264 0.1729 1.0000
14.500 1.4181 0.04872 0.04068 -0.0253 0.1692 1.0000
14.750 1.4200 0.05090 0.04289 -0.0243 0.1661 1.0000
15.000 1.4207 0.05322 0.04523 -0.0234 0.1632 1.0000
15.250 1.4259 0.05519 0.04730 -0.0227 0.1601 1.0000
15.500 1.4294 0.05736 0.04956 -0.0221 0.1569 1.0000
15.750 1.4311 0.05974 0.05200 -0.0215 0.1538 1.0000
16.000 1.4313 0.06230 0.05459 -0.0210 0.1509 1.0000
16.250 1.4314 0.06488 0.05720 -0.0206 0.1481 1.0000
16.500 1.4348 0.06723 0.05970 -0.0203 0.1451 1.0000
16.750 1.4356 0.06988 0.06244 -0.0201 0.1416 1.0000
17.000 1.4345 0.07276 0.06538 -0.0200 0.1383 1.0000
17.250 1.4313 0.07589 0.06852 -0.0200 0.1354 1.0000
17.500 1.4329 0.07857 0.07135 -0.0201 0.1320 1.0000
17.750 1.4312 0.08169 0.07457 -0.0203 0.1281 1.0000
18.000 1.4260 0.08525 0.07818 -0.0207 0.1242 1.0000
18.250 1.4243 0.08842 0.08143 -0.0210 0.1209 1.0000
18.500 1.4211 0.09185 0.08496 -0.0215 0.1168 1.0000
18.750 1.4149 0.09568 0.08883 -0.0222 0.1132 1.0000
19.000 1.4112 0.09917 0.09239 -0.0229 0.1102 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 481A AIRFOIL (goe481a-il)