GOE 100 (SOPWITH) AIRFOIL (goe100-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: GOE 100 (SOPWITH) AIRFOIL (goe100-il) Reynolds number: 100,000 Max Cl/Cd: 54.41 at α=3.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe100-il-100000.txt Download as CSV file: xf-goe100-il-100000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 100 (SOPWITH) 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
------ -------- --------- --------- -------- -------- --------
-9.000 -0.3112 0.10194 0.09737 -0.0227 1.0000 0.0612
-8.750 -0.3093 0.09869 0.09416 -0.0230 1.0000 0.0629
-8.500 -0.4097 0.10696 0.10218 -0.0190 1.0000 0.0579
-8.250 -0.4016 0.10315 0.09839 -0.0184 1.0000 0.0596
-8.000 -0.3974 0.10002 0.09530 -0.0187 1.0000 0.0612
-7.750 -0.3954 0.09712 0.09245 -0.0192 1.0000 0.0629
-7.500 -0.3961 0.09445 0.08985 -0.0196 1.0000 0.0646
-7.250 -0.3980 0.09187 0.08736 -0.0206 1.0000 0.0663
-7.000 -0.3983 0.08946 0.08501 -0.0242 1.0000 0.0682
-6.750 -0.3955 0.08773 0.08328 -0.0314 1.0000 0.0694
-6.500 -0.3898 0.08441 0.07996 -0.0346 1.0000 0.0700
-6.250 -0.3877 0.07924 0.07491 -0.0291 1.0000 0.0714
-6.000 -0.3828 0.07606 0.07177 -0.0270 1.0000 0.0734
-5.750 -0.3769 0.07313 0.06885 -0.0267 1.0000 0.0760
-5.500 -0.3685 0.07017 0.06582 -0.0282 1.0000 0.0798
-5.250 -0.3511 0.06689 0.06233 -0.0346 1.0000 0.0838
-5.000 -0.3488 0.06341 0.05898 -0.0306 1.0000 0.0863
-4.750 -0.3235 0.06149 0.05667 -0.0360 1.0000 0.0967
-4.500 -0.3233 0.05738 0.05284 -0.0315 1.0000 0.1001
-4.250 -0.3012 0.05449 0.04969 -0.0344 1.0000 0.1110
-4.000 -0.2912 0.05142 0.04670 -0.0323 1.0000 0.1154
-3.750 -0.2699 0.04834 0.04343 -0.0339 1.0000 0.1262
-3.500 -0.2478 0.04562 0.04050 -0.0349 1.0000 0.1389
-3.250 -0.2263 0.04300 0.03771 -0.0354 1.0000 0.1523
-3.000 -0.2041 0.04035 0.03490 -0.0357 1.0000 0.1660
-2.750 -0.1423 0.02937 0.02237 -0.0399 1.0000 0.0670
-2.500 -0.1106 0.02681 0.01903 -0.0401 1.0000 0.0696
-2.250 -0.0776 0.02363 0.01535 -0.0410 1.0000 0.0718
-2.000 -0.0473 0.02185 0.01329 -0.0412 1.0000 0.0781
-1.750 -0.0182 0.02071 0.01181 -0.0411 1.0000 0.0969
-1.500 0.0085 0.02006 0.01112 -0.0410 1.0000 0.1224
-1.250 0.0329 0.01982 0.01090 -0.0409 1.0000 0.1465
-1.000 0.0599 0.01920 0.01021 -0.0408 1.0000 0.1589
-0.750 0.0859 0.01885 0.00977 -0.0407 1.0000 0.1773
-0.500 0.1130 0.01829 0.00936 -0.0409 1.0000 0.2039
-0.250 0.1436 0.01745 0.00895 -0.0417 1.0000 0.2865
0.000 0.1639 0.01552 0.00888 -0.0400 1.0000 1.0000
0.250 0.1960 0.01592 0.00894 -0.0414 0.9967 1.0000
0.500 0.2522 0.01627 0.00902 -0.0475 0.9847 1.0000
0.750 0.3077 0.01645 0.00902 -0.0532 0.9712 1.0000
1.000 0.3636 0.01643 0.00888 -0.0588 0.9567 1.0000
1.250 0.4124 0.01623 0.00863 -0.0626 0.9396 1.0000
1.500 0.4624 0.01589 0.00828 -0.0664 0.9230 1.0000
1.750 0.5135 0.01539 0.00781 -0.0701 0.9073 1.0000
2.000 0.5596 0.01483 0.00728 -0.0726 0.8903 1.0000
2.250 0.5956 0.01438 0.00687 -0.0729 0.8680 1.0000
2.500 0.6288 0.01393 0.00646 -0.0726 0.8419 1.0000
2.750 0.6582 0.01357 0.00616 -0.0715 0.8088 1.0000
3.000 0.6846 0.01331 0.00586 -0.0698 0.7633 1.0000
3.250 0.7096 0.01319 0.00558 -0.0679 0.6957 1.0000
3.500 0.7329 0.01347 0.00539 -0.0657 0.6024 1.0000
3.750 0.7532 0.01426 0.00562 -0.0637 0.5199 1.0000
4.000 0.7724 0.01511 0.00605 -0.0619 0.4495 1.0000
4.250 0.7924 0.01584 0.00642 -0.0604 0.3938 1.0000
4.500 0.8134 0.01648 0.00681 -0.0593 0.3498 1.0000
4.750 0.8358 0.01705 0.00725 -0.0584 0.3173 1.0000
5.000 0.8588 0.01759 0.00772 -0.0576 0.2930 1.0000
5.250 0.8812 0.01802 0.00812 -0.0568 0.2605 1.0000
5.500 0.9038 0.01850 0.00846 -0.0560 0.2286 1.0000
5.750 0.9278 0.01896 0.00896 -0.0553 0.2087 1.0000
6.000 0.9513 0.01946 0.00943 -0.0546 0.1827 1.0000
6.250 0.9753 0.01995 0.00993 -0.0538 0.1622 1.0000
6.500 0.9988 0.02055 0.01045 -0.0530 0.1278 1.0000
6.750 1.0150 0.02233 0.01162 -0.0514 0.0444 1.0000
7.000 1.0355 0.02356 0.01292 -0.0500 0.0384 1.0000
7.250 1.0569 0.02461 0.01422 -0.0486 0.0371 1.0000
7.500 1.0772 0.02577 0.01564 -0.0472 0.0364 1.0000
7.750 1.0957 0.02704 0.01719 -0.0455 0.0362 1.0000
8.000 1.1121 0.02844 0.01883 -0.0435 0.0362 1.0000
8.250 1.1264 0.02997 0.02058 -0.0412 0.0365 1.0000
8.500 1.1400 0.03166 0.02242 -0.0389 0.0370 1.0000
8.750 1.1549 0.03361 0.02454 -0.0367 0.0376 1.0000
9.000 1.1737 0.03593 0.02692 -0.0351 0.0381 1.0000
9.250 1.1945 0.03838 0.02953 -0.0338 0.0379 1.0000
9.500 1.2149 0.04118 0.03256 -0.0325 0.0378 1.0000
9.750 1.2390 0.04544 0.03701 -0.0320 0.0391 1.0000
10.000 1.2530 0.04641 0.03855 -0.0291 0.0413 1.0000
10.250 1.2575 0.05019 0.04302 -0.0260 0.0443 1.0000
10.500 1.2569 0.05456 0.04787 -0.0231 0.0467 1.0000
10.750 1.2561 0.05984 0.05344 -0.0210 0.0487 1.0000
11.000 1.2533 0.06261 0.05667 -0.0178 0.0511 1.0000
11.250 1.2276 0.06604 0.06050 -0.0137 0.0519 1.0000
11.500 1.2013 0.07008 0.06485 -0.0115 0.0522 1.0000
11.750 1.1757 0.07480 0.06982 -0.0112 0.0523 1.0000
12.000 1.1510 0.08025 0.07547 -0.0127 0.0523 1.0000
12.250 1.1275 0.08645 0.08184 -0.0157 0.0520 1.0000
12.500 1.1046 0.09367 0.08918 -0.0203 0.0517 1.0000
12.750 1.0828 0.10174 0.09734 -0.0259 0.0513 1.0000
13.000 1.0600 0.11087 0.10653 -0.0323 0.0507 1.0000
13.250 1.0359 0.12116 0.11683 -0.0393 0.0502 1.0000
13.500 0.9823 0.14317 0.13876 -0.0545 0.0516 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 100 (SOPWITH) AIRFOIL (goe100-il)