GOE 624 AIRFOIL (goe624-il) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: GOE 624 AIRFOIL (goe624-il) Reynolds number: 200,000 Max Cl/Cd: 69.05 at α=7° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe624-il-200000.txt Download as CSV file: xf-goe624-il-200000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 624 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
------ -------- --------- --------- -------- -------- --------
-10.000 -0.0622 0.09168 0.08827 -0.0784 0.9421 0.0834
-9.750 -0.0656 0.08807 0.08465 -0.0801 0.9322 0.0864
-9.500 -0.1109 0.08069 0.07728 -0.0885 0.9226 0.0886
-9.250 -0.0908 0.07846 0.07503 -0.0847 0.9116 0.0893
-9.000 -0.0801 0.07633 0.07287 -0.0824 0.9010 0.0901
-8.750 -0.0731 0.07398 0.07048 -0.0811 0.8895 0.0912
-8.500 -0.0693 0.07125 0.06773 -0.0807 0.8790 0.0926
-8.250 -0.3022 0.04119 0.03611 -0.1104 0.8908 0.0618
-8.000 -0.2915 0.03794 0.03264 -0.1095 0.8828 0.0615
-7.750 -0.2824 0.03458 0.02884 -0.1084 0.8729 0.0615
-7.500 -0.2700 0.03170 0.02544 -0.1070 0.8643 0.0618
-7.250 -0.2533 0.02927 0.02269 -0.1058 0.8546 0.0623
-7.000 -0.2328 0.02751 0.02072 -0.1049 0.8467 0.0628
-6.750 -0.2105 0.02618 0.01923 -0.1041 0.8377 0.0636
-6.500 -0.1865 0.02503 0.01789 -0.1034 0.8317 0.0645
-6.250 -0.1626 0.02395 0.01665 -0.1029 0.8227 0.0656
-6.000 -0.1376 0.02290 0.01535 -0.1022 0.8160 0.0672
-5.750 -0.1122 0.02203 0.01418 -0.1016 0.8087 0.0692
-5.500 -0.0867 0.02104 0.01312 -0.1012 0.8010 0.0711
-5.250 -0.0599 0.02041 0.01242 -0.1008 0.7952 0.0733
-5.000 -0.0336 0.01980 0.01172 -0.1004 0.7869 0.0761
-4.750 -0.0067 0.01904 0.01080 -0.1000 0.7802 0.0793
-4.500 0.0201 0.01856 0.01036 -0.0997 0.7736 0.0835
-4.250 0.0468 0.01804 0.00975 -0.0993 0.7656 0.0897
-4.000 0.0743 0.01761 0.00929 -0.0990 0.7597 0.0975
-3.750 0.1006 0.01711 0.00885 -0.0987 0.7518 0.1070
-3.500 0.1277 0.01667 0.00839 -0.0985 0.7450 0.1177
-3.250 0.1552 0.01636 0.00803 -0.0982 0.7386 0.1285
-3.000 0.1822 0.01612 0.00780 -0.0980 0.7303 0.1384
-2.750 0.2099 0.01576 0.00742 -0.0978 0.7240 0.1480
-2.500 0.2371 0.01559 0.00724 -0.0976 0.7156 0.1584
-2.250 0.2648 0.01536 0.00701 -0.0975 0.7086 0.1719
-2.000 0.2920 0.01513 0.00683 -0.0973 0.7013 0.1871
-1.750 0.3192 0.01484 0.00660 -0.0971 0.6932 0.2048
-1.500 0.3465 0.01449 0.00634 -0.0969 0.6863 0.2336
-1.250 0.3722 0.01405 0.00620 -0.0966 0.6772 0.2952
-1.000 0.3976 0.01346 0.00602 -0.0962 0.6705 0.4218
-0.750 0.4174 0.01273 0.00612 -0.0945 0.6613 0.6286
-0.500 0.4371 0.01225 0.00621 -0.0915 0.6543 0.8390
-0.250 0.4823 0.01227 0.00628 -0.0940 0.6447 0.9497
0.000 0.5437 0.01235 0.00613 -0.1005 0.6358 0.9912
0.250 0.5842 0.01243 0.00611 -0.1033 0.6254 1.0000
0.500 0.6088 0.01253 0.00601 -0.1026 0.6179 1.0000
0.750 0.6316 0.01265 0.00608 -0.1018 0.6085 1.0000
1.000 0.6565 0.01277 0.00603 -0.1012 0.6010 1.0000
1.250 0.6797 0.01292 0.00613 -0.1004 0.5919 1.0000
1.500 0.7046 0.01306 0.00612 -0.0997 0.5843 1.0000
1.750 0.7286 0.01324 0.00624 -0.0991 0.5763 1.0000
2.000 0.7535 0.01340 0.00630 -0.0985 0.5688 1.0000
2.250 0.7788 0.01359 0.00639 -0.0980 0.5620 1.0000
2.500 0.8034 0.01377 0.00653 -0.0975 0.5543 1.0000
2.750 0.8300 0.01396 0.00658 -0.0972 0.5484 1.0000
3.000 0.8545 0.01418 0.00680 -0.0967 0.5413 1.0000
3.250 0.8803 0.01437 0.00693 -0.0964 0.5349 1.0000
3.500 0.9073 0.01460 0.00705 -0.0963 0.5295 1.0000
3.750 0.9318 0.01483 0.00731 -0.0958 0.5228 1.0000
4.000 0.9581 0.01504 0.00745 -0.0955 0.5171 1.0000
4.250 0.9849 0.01531 0.00763 -0.0954 0.5119 1.0000
4.500 1.0094 0.01555 0.00792 -0.0950 0.5056 1.0000
4.750 1.0357 0.01577 0.00808 -0.0948 0.4999 1.0000
5.000 1.0614 0.01602 0.00828 -0.0945 0.4936 1.0000
5.250 1.0854 0.01622 0.00849 -0.0939 0.4862 1.0000
5.500 1.1125 0.01646 0.00861 -0.0939 0.4802 1.0000
5.750 1.1362 0.01676 0.00897 -0.0933 0.4743 1.0000
6.000 1.1612 0.01704 0.00925 -0.0930 0.4687 1.0000
6.250 1.1883 0.01733 0.00947 -0.0930 0.4638 1.0000
6.500 1.2120 0.01768 0.00987 -0.0925 0.4583 1.0000
6.750 1.2359 0.01797 0.01020 -0.0920 0.4525 1.0000
7.000 1.2623 0.01828 0.01045 -0.0919 0.4475 1.0000
7.250 1.2864 0.01867 0.01088 -0.0916 0.4424 1.0000
7.500 1.3089 0.01901 0.01129 -0.0909 0.4367 1.0000
7.750 1.3343 0.01933 0.01158 -0.0907 0.4316 1.0000
8.000 1.3590 0.01974 0.01199 -0.0904 0.4264 1.0000
8.250 1.3791 0.02010 0.01245 -0.0893 0.4202 1.0000
8.500 1.4034 0.02041 0.01274 -0.0890 0.4144 1.0000
8.750 1.4260 0.02083 0.01319 -0.0884 0.4087 1.0000
9.000 1.4446 0.02121 0.01367 -0.0871 0.4023 1.0000
9.250 1.4684 0.02153 0.01395 -0.0867 0.3965 1.0000
9.500 1.4858 0.02198 0.01449 -0.0852 0.3899 1.0000
9.750 1.5029 0.02232 0.01489 -0.0837 0.3828 1.0000
10.000 1.5244 0.02269 0.01522 -0.0830 0.3763 1.0000
10.250 1.5340 0.02314 0.01582 -0.0803 0.3688 1.0000
10.500 1.5530 0.02347 0.01608 -0.0791 0.3619 1.0000
10.750 1.5562 0.02403 0.01680 -0.0755 0.3545 1.0000
11.000 1.5674 0.02451 0.01730 -0.0732 0.3476 1.0000
11.250 1.5764 0.02518 0.01804 -0.0709 0.3405 1.0000
11.500 1.5832 0.02590 0.01882 -0.0683 0.3329 1.0000
11.750 1.5918 0.02669 0.01964 -0.0662 0.3252 1.0000
12.000 1.5965 0.02765 0.02066 -0.0638 0.3170 1.0000
12.250 1.6044 0.02863 0.02168 -0.0619 0.3095 1.0000
12.500 1.6098 0.02979 0.02291 -0.0599 0.3017 1.0000
12.750 1.6178 0.03093 0.02405 -0.0583 0.2946 1.0000
13.000 1.6218 0.03233 0.02553 -0.0565 0.2869 1.0000
13.250 1.6296 0.03358 0.02676 -0.0551 0.2801 1.0000
13.500 1.6319 0.03524 0.02854 -0.0534 0.2727 1.0000
13.750 1.6401 0.03653 0.02975 -0.0521 0.2661 1.0000
14.000 1.6398 0.03853 0.03190 -0.0506 0.2590 1.0000
14.250 1.6465 0.03998 0.03328 -0.0494 0.2525 1.0000
14.500 1.6448 0.04221 0.03567 -0.0481 0.2457 1.0000
14.750 1.6469 0.04413 0.03758 -0.0470 0.2391 1.0000
15.000 1.6464 0.04641 0.03994 -0.0459 0.2326 1.0000
15.250 1.6451 0.04884 0.04243 -0.0450 0.2259 1.0000
15.500 1.6469 0.05101 0.04460 -0.0442 0.2198 1.0000
15.750 1.6424 0.05397 0.04770 -0.0436 0.2135 1.0000
16.000 1.6460 0.05596 0.04959 -0.0429 0.2072 1.0000
16.250 1.6377 0.05958 0.05342 -0.0427 0.2011 1.0000
16.500 1.6355 0.06242 0.05626 -0.0423 0.1948 1.0000
16.750 1.6296 0.06587 0.05981 -0.0422 0.1886 1.0000
17.000 1.6229 0.06946 0.06346 -0.0422 0.1823 1.0000
17.250 1.6183 0.07279 0.06682 -0.0422 0.1762 1.0000
17.500 1.6092 0.07693 0.07110 -0.0426 0.1700 1.0000
17.750 1.6074 0.07992 0.07402 -0.0426 0.1643 1.0000
18.000 1.5959 0.08457 0.07888 -0.0434 0.1585 1.0000
18.250 1.5907 0.08821 0.08251 -0.0439 0.1530 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 624 AIRFOIL (goe624-il)