GOE 506 AIRFOIL (goe506-il) Xfoil prediction polar at RE=50,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: GOE 506 AIRFOIL (goe506-il) Reynolds number: 50,000 Max Cl/Cd: 27.98 at α=7.5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe506-il-50000-n5.txt Download as CSV file: xf-goe506-il-50000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 506 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.050 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.000 -0.3957 0.10724 0.10110 -0.0509 0.9875 0.0583
-9.750 -0.4056 0.10056 0.09446 -0.0551 0.9840 0.0580
-9.500 -0.4256 0.09260 0.08655 -0.0598 0.9791 0.0571
-9.250 -0.4597 0.08239 0.07627 -0.0681 0.9738 0.0557
-9.000 -0.5108 0.07584 0.06958 -0.0698 0.9662 0.0542
-8.750 -0.5425 0.07022 0.06363 -0.0704 0.9606 0.0534
-8.500 -0.5565 0.06665 0.05983 -0.0689 0.9550 0.0538
-8.250 -0.5616 0.06288 0.05572 -0.0683 0.9505 0.0541
-8.000 -0.5622 0.05961 0.05210 -0.0672 0.9469 0.0547
-7.750 -0.5638 0.05684 0.04900 -0.0648 0.9429 0.0551
-7.500 -0.5575 0.05424 0.04604 -0.0634 0.9392 0.0560
-7.250 -0.5452 0.05188 0.04328 -0.0626 0.9359 0.0577
-7.000 -0.5287 0.04969 0.04056 -0.0621 0.9332 0.0599
-6.750 -0.5207 0.04807 0.03849 -0.0595 0.9298 0.0612
-6.500 -0.5060 0.04637 0.03673 -0.0581 0.9266 0.0627
-6.250 -0.4876 0.04501 0.03524 -0.0572 0.9233 0.0646
-6.000 -0.4664 0.04383 0.03388 -0.0566 0.9205 0.0677
-5.750 -0.4421 0.04279 0.03254 -0.0564 0.9180 0.0714
-5.500 -0.4272 0.04186 0.03143 -0.0544 0.9140 0.0734
-5.250 -0.4106 0.04085 0.03039 -0.0530 0.9098 0.0761
-5.000 -0.3868 0.04001 0.02944 -0.0527 0.9054 0.0813
-4.750 -0.3581 0.03924 0.02847 -0.0531 0.9001 0.0872
-4.500 -0.3391 0.03836 0.02757 -0.0518 0.8923 0.0923
-4.250 -0.3063 0.03754 0.02671 -0.0530 0.8869 0.1035
-4.000 -0.2867 0.03672 0.02615 -0.0521 0.8798 0.1252
-3.750 -0.2589 0.03637 0.02614 -0.0524 0.8739 0.2053
-3.500 -0.2313 0.03649 0.02620 -0.0525 0.8681 0.2613
-3.250 -0.2123 0.03659 0.02622 -0.0511 0.8601 0.2949
-3.000 -0.1803 0.03674 0.02629 -0.0519 0.8538 0.3319
-2.750 -0.1597 0.03666 0.02620 -0.0506 0.8433 0.3565
-2.500 -0.1319 0.03648 0.02590 -0.0504 0.8326 0.3814
-2.250 -0.0895 0.03604 0.02537 -0.0524 0.8235 0.4055
-2.000 -0.0596 0.03558 0.02477 -0.0525 0.8113 0.4195
-1.750 -0.0304 0.03514 0.02422 -0.0525 0.7995 0.4319
-1.500 0.0049 0.03467 0.02369 -0.0535 0.7909 0.4486
-1.250 0.0350 0.03426 0.02328 -0.0537 0.7814 0.4693
-1.000 0.0591 0.03390 0.02300 -0.0529 0.7706 0.4929
-0.750 0.0971 0.03320 0.02244 -0.0541 0.7638 0.5328
-0.500 0.1166 0.03280 0.02228 -0.0524 0.7515 0.5832
-0.250 0.1437 0.03215 0.02213 -0.0515 0.7407 0.6851
0.000 0.2790 0.03170 0.02201 -0.0710 0.7349 1.0000
0.250 0.2974 0.03154 0.02163 -0.0693 0.7214 1.0000
0.500 0.3183 0.03132 0.02121 -0.0679 0.7083 1.0000
0.750 0.3519 0.03071 0.02038 -0.0681 0.6997 1.0000
1.000 0.3698 0.03055 0.02007 -0.0662 0.6846 1.0000
1.250 0.3879 0.03041 0.01978 -0.0642 0.6688 1.0000
1.750 0.4175 0.03050 0.01962 -0.0597 0.6306 1.0000
2.000 0.4390 0.03037 0.01935 -0.0584 0.6136 1.0000
2.250 0.4646 0.03012 0.01895 -0.0575 0.5981 1.0000
2.500 0.4938 0.02976 0.01843 -0.0571 0.5840 1.0000
2.750 0.5266 0.02933 0.01780 -0.0573 0.5718 1.0000
3.000 0.5631 0.02880 0.01704 -0.0578 0.5605 1.0000
3.250 0.5942 0.02863 0.01667 -0.0579 0.5474 1.0000
3.500 0.6244 0.02860 0.01644 -0.0580 0.5352 1.0000
3.750 0.6564 0.02859 0.01622 -0.0584 0.5244 1.0000
4.000 0.6877 0.02871 0.01615 -0.0588 0.5142 1.0000
4.250 0.7129 0.02910 0.01643 -0.0585 0.5042 1.0000
4.500 0.7475 0.02925 0.01638 -0.0594 0.4958 1.0000
4.750 0.7667 0.02988 0.01699 -0.0584 0.4868 1.0000
5.000 0.8006 0.03013 0.01710 -0.0594 0.4793 1.0000
5.250 0.8171 0.03084 0.01782 -0.0580 0.4712 1.0000
5.500 0.8464 0.03124 0.01815 -0.0584 0.4644 1.0000
5.750 0.8655 0.03189 0.01880 -0.0574 0.4569 1.0000
6.000 0.8877 0.03246 0.01935 -0.0568 0.4499 1.0000
6.250 0.9139 0.03295 0.01981 -0.0568 0.4436 1.0000
6.500 0.9277 0.03376 0.02070 -0.0551 0.4373 1.0000
6.750 0.9553 0.03422 0.02113 -0.0553 0.4317 1.0000
7.000 0.9720 0.03499 0.02197 -0.0541 0.4260 1.0000
7.250 0.9858 0.03584 0.02290 -0.0525 0.4203 1.0000
7.500 1.0143 0.03625 0.02328 -0.0528 0.4148 1.0000
7.750 1.0229 0.03729 0.02443 -0.0505 0.4094 1.0000
8.000 1.0347 0.03824 0.02549 -0.0487 0.4039 1.0000
8.250 1.0627 0.03868 0.02594 -0.0490 0.3991 1.0000
8.500 1.0678 0.03990 0.02729 -0.0464 0.3938 1.0000
8.750 1.0744 0.04105 0.02855 -0.0441 0.3879 1.0000
9.000 1.1031 0.04140 0.02892 -0.0443 0.3829 1.0000
9.250 1.1013 0.04300 0.03069 -0.0413 0.3777 1.0000
9.500 1.1027 0.04458 0.03242 -0.0387 0.3725 1.0000
9.750 1.1257 0.04519 0.03310 -0.0384 0.3680 1.0000
10.000 1.1350 0.04649 0.03450 -0.0367 0.3633 1.0000
10.250 1.1146 0.04936 0.03761 -0.0326 0.3575 1.0000
10.500 1.1297 0.05036 0.03870 -0.0316 0.3525 1.0000
10.750 1.1536 0.05088 0.03927 -0.0313 0.3478 1.0000
11.000 1.0961 0.05667 0.04533 -0.0256 0.3404 1.0000
11.250 1.1137 0.05747 0.04621 -0.0248 0.3352 1.0000
11.500 1.1160 0.05951 0.04835 -0.0233 0.3297 1.0000
11.750 1.0447 0.06829 0.05732 -0.0200 0.3192 1.0000
12.000 1.0877 0.06659 0.05568 -0.0197 0.3162 1.0000
12.250 0.9886 0.08027 0.06948 -0.0187 0.3009 1.0000
12.500 1.0202 0.07934 0.06864 -0.0179 0.2988 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 506 AIRFOIL (goe506-il)