GOE 402 AIRFOIL (goe402-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: GOE 402 AIRFOIL (goe402-il) Reynolds number: 100,000 Max Cl/Cd: 55.07 at α=9.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe402-il-100000.txt Download as CSV file: xf-goe402-il-100000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 402 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
------ -------- --------- --------- -------- -------- --------
-7.500 -0.2794 0.09509 0.09051 -0.0218 1.0000 0.0420
-7.250 -0.2794 0.09296 0.08847 -0.0215 1.0000 0.0428
-7.000 -0.2796 0.09084 0.08644 -0.0214 1.0000 0.0437
-6.750 -0.2767 0.08859 0.08427 -0.0222 1.0000 0.0446
-6.500 -0.2732 0.08665 0.08241 -0.0240 1.0000 0.0460
-6.250 -0.2675 0.08532 0.08115 -0.0275 1.0000 0.0469
-6.000 -0.2577 0.08499 0.08084 -0.0333 1.0000 0.0475
-5.750 -0.2536 0.08413 0.08000 -0.0351 1.0000 0.0477
-5.500 -0.2664 0.08061 0.07665 -0.0307 1.0000 0.0482
-5.250 -0.2798 0.07819 0.07435 -0.0258 1.0000 0.0487
-5.000 -0.2673 0.07403 0.07024 -0.0257 0.9957 0.0506
-4.750 -0.2274 0.06942 0.06556 -0.0334 0.9868 0.0539
-4.500 -0.1476 0.06680 0.06248 -0.0526 0.9751 0.0598
-4.250 -0.1246 0.06019 0.05597 -0.0552 0.9677 0.0617
-4.000 -0.0911 0.05604 0.05179 -0.0593 0.9585 0.0662
-3.750 -0.0332 0.05225 0.04767 -0.0697 0.9475 0.0748
-3.500 0.0193 0.04924 0.04434 -0.0772 0.9378 0.0881
-3.250 0.0531 0.04428 0.03949 -0.0804 0.9291 0.0967
-3.000 0.0981 0.04089 0.03589 -0.0855 0.9170 0.1111
-2.750 0.1422 0.03793 0.03265 -0.0901 0.9047 0.1328
-1.500 0.2789 0.02595 0.02046 -0.0925 0.8418 0.3940
-1.250 0.2990 0.02373 0.01827 -0.0905 0.8279 0.4548
-1.000 0.3311 0.02237 0.01665 -0.0910 0.8137 0.4795
-0.750 0.4191 0.02235 0.01446 -0.0979 0.8002 0.1438
-0.500 0.4488 0.02102 0.01275 -0.0971 0.7861 0.1270
-0.250 0.4780 0.02019 0.01148 -0.0960 0.7719 0.1147
0.000 0.5049 0.01932 0.01042 -0.0951 0.7576 0.1106
0.250 0.5316 0.01874 0.00967 -0.0941 0.7432 0.1100
0.500 0.5577 0.01835 0.00912 -0.0931 0.7286 0.1133
0.750 0.5835 0.01793 0.00860 -0.0920 0.7137 0.1134
1.000 0.6086 0.01759 0.00818 -0.0908 0.6986 0.1147
1.250 0.6338 0.01732 0.00781 -0.0897 0.6832 0.1178
1.500 0.6594 0.01715 0.00755 -0.0887 0.6676 0.1249
1.750 0.6852 0.01703 0.00737 -0.0878 0.6518 0.1446
2.000 0.7191 0.01526 0.00723 -0.0886 0.6355 1.0000
2.250 0.7440 0.01547 0.00714 -0.0875 0.6179 1.0000
2.500 0.7687 0.01569 0.00714 -0.0865 0.6005 1.0000
2.750 0.7937 0.01594 0.00720 -0.0856 0.5841 1.0000
3.000 0.8189 0.01624 0.00734 -0.0849 0.5689 1.0000
3.250 0.8440 0.01657 0.00753 -0.0842 0.5545 1.0000
3.500 0.8692 0.01693 0.00777 -0.0835 0.5411 1.0000
3.750 0.8946 0.01731 0.00804 -0.0830 0.5291 1.0000
4.000 0.9207 0.01770 0.00832 -0.0825 0.5190 1.0000
4.250 0.9452 0.01816 0.00880 -0.0819 0.5076 1.0000
4.500 0.9701 0.01861 0.00924 -0.0814 0.4973 1.0000
4.750 0.9957 0.01900 0.00955 -0.0809 0.4880 1.0000
5.000 1.0203 0.01940 0.01001 -0.0803 0.4779 1.0000
5.250 1.0447 0.01989 0.01055 -0.0798 0.4689 1.0000
5.500 1.0707 0.02033 0.01095 -0.0794 0.4617 1.0000
5.750 1.0943 0.02090 0.01167 -0.0788 0.4533 1.0000
6.000 1.1207 0.02132 0.01207 -0.0785 0.4465 1.0000
6.250 1.1433 0.02195 0.01291 -0.0778 0.4385 1.0000
6.500 1.1698 0.02240 0.01334 -0.0775 0.4321 1.0000
6.750 1.1917 0.02299 0.01416 -0.0767 0.4232 1.0000
7.000 1.2175 0.02348 0.01467 -0.0763 0.4165 1.0000
7.250 1.2394 0.02415 0.01562 -0.0755 0.4085 1.0000
7.500 1.2647 0.02469 0.01623 -0.0750 0.4016 1.0000
7.750 1.2867 0.02531 0.01708 -0.0742 0.3928 1.0000
8.000 1.3094 0.02591 0.01786 -0.0734 0.3839 1.0000
8.250 1.3319 0.02607 0.01813 -0.0723 0.3705 1.0000
8.500 1.3517 0.02588 0.01799 -0.0707 0.3513 1.0000
8.750 1.3712 0.02559 0.01761 -0.0690 0.3301 1.0000
9.000 1.3850 0.02555 0.01778 -0.0667 0.3080 1.0000
9.250 1.3993 0.02555 0.01794 -0.0645 0.2865 1.0000
9.500 1.4086 0.02558 0.01829 -0.0616 0.2535 1.0000
9.750 1.4030 0.02684 0.01907 -0.0572 0.1541 1.0000
10.000 1.3844 0.03033 0.02193 -0.0523 0.0975 1.0000
10.250 1.3738 0.03286 0.02426 -0.0481 0.0713 1.0000
10.500 1.3616 0.03551 0.02679 -0.0445 0.0633 1.0000
10.750 1.3530 0.03814 0.02958 -0.0419 0.0589 1.0000
11.000 1.3429 0.04122 0.03278 -0.0401 0.0560 1.0000
11.250 1.3315 0.04479 0.03642 -0.0391 0.0539 1.0000
11.500 1.3219 0.04839 0.04007 -0.0383 0.0521 1.0000
11.750 1.3210 0.05130 0.04313 -0.0374 0.0500 1.0000
12.000 1.3223 0.05400 0.04593 -0.0363 0.0475 1.0000
12.250 1.3250 0.05665 0.04859 -0.0353 0.0447 1.0000
12.500 1.3458 0.05825 0.05005 -0.0321 0.0413 1.0000
12.750 1.3604 0.06054 0.05257 -0.0307 0.0403 1.0000
13.000 1.3739 0.06329 0.05557 -0.0295 0.0395 1.0000
13.250 1.3815 0.06670 0.05926 -0.0286 0.0390 1.0000
13.500 1.3823 0.07059 0.06344 -0.0282 0.0388 1.0000
13.750 1.3763 0.07492 0.06806 -0.0284 0.0384 1.0000
14.000 1.3668 0.07960 0.07302 -0.0291 0.0382 1.0000
14.250 1.3545 0.08472 0.07841 -0.0304 0.0381 1.0000
14.500 1.3390 0.09032 0.08427 -0.0324 0.0380 1.0000
14.750 1.3220 0.09638 0.09059 -0.0349 0.0381 1.0000
15.000 1.3034 0.10303 0.09747 -0.0381 0.0385 1.0000
15.250 1.2836 0.11019 0.10486 -0.0420 0.0389 1.0000
15.500 1.2629 0.11802 0.11289 -0.0466 0.0393 1.0000
15.750 1.2421 0.12640 0.12144 -0.0518 0.0398 1.0000
16.000 1.2221 0.13523 0.13042 -0.0573 0.0405 1.0000
16.250 1.2034 0.14434 0.13964 -0.0629 0.0411 1.0000
16.500 1.1882 0.15323 0.14864 -0.0683 0.0418 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 402 AIRFOIL (goe402-il)