GOE 702 AIRFOIL (goe702-il) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: GOE 702 AIRFOIL (goe702-il) Reynolds number: 200,000 Max Cl/Cd: 57.12 at α=6.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe702-il-200000.txt Download as CSV file: xf-goe702-il-200000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 702 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 -0.1275 0.09798 0.09363 -0.0552 0.7869 0.1020
-9.750 -0.3154 0.06196 0.05762 -0.0798 0.7898 0.0788
-9.500 -0.3737 0.05160 0.04720 -0.0863 0.7831 0.0776
-9.250 -0.4528 0.04059 0.03549 -0.0878 0.7771 0.0769
-9.000 -0.4700 0.03552 0.02973 -0.0861 0.7718 0.0778
-8.750 -0.4675 0.03246 0.02611 -0.0845 0.7668 0.0791
-8.500 -0.4449 0.03155 0.02526 -0.0841 0.7605 0.0803
-8.250 -0.4229 0.03078 0.02444 -0.0833 0.7551 0.0815
-8.000 -0.4036 0.02961 0.02307 -0.0823 0.7504 0.0830
-7.750 -0.3856 0.02815 0.02133 -0.0813 0.7450 0.0848
-7.500 -0.3680 0.02678 0.01949 -0.0800 0.7391 0.0870
-7.250 -0.3440 0.02584 0.01860 -0.0795 0.7339 0.0888
-7.000 -0.3196 0.02534 0.01803 -0.0788 0.7293 0.0908
-6.750 -0.2959 0.02462 0.01718 -0.0781 0.7235 0.0933
-6.500 -0.2731 0.02379 0.01602 -0.0772 0.7173 0.0959
-6.250 -0.2482 0.02294 0.01518 -0.0766 0.7120 0.0984
-6.000 -0.2227 0.02257 0.01474 -0.0759 0.7070 0.1014
-5.750 -0.1976 0.02205 0.01408 -0.0753 0.7004 0.1049
-5.500 -0.1724 0.02125 0.01319 -0.0747 0.6944 0.1081
-5.250 -0.1462 0.02085 0.01276 -0.0741 0.6891 0.1118
-5.000 -0.1204 0.02050 0.01227 -0.0735 0.6822 0.1163
-4.750 -0.0946 0.01980 0.01158 -0.0730 0.6753 0.1205
-4.500 -0.0679 0.01944 0.01115 -0.0724 0.6697 0.1254
-4.250 -0.0417 0.01903 0.01061 -0.0718 0.6628 0.1309
-4.000 -0.0153 0.01854 0.01021 -0.0713 0.6551 0.1363
-3.750 0.0118 0.01826 0.00974 -0.0707 0.6489 0.1426
-3.500 0.0379 0.01773 0.00931 -0.0703 0.6405 0.1487
-3.250 0.0650 0.01744 0.00894 -0.0697 0.6327 0.1554
-3.000 0.0919 0.01702 0.00848 -0.0692 0.6259 0.1621
-2.750 0.1185 0.01675 0.00824 -0.0687 0.6162 0.1691
-2.500 0.1459 0.01639 0.00778 -0.0682 0.6089 0.1764
-2.250 0.1723 0.01614 0.00760 -0.0677 0.5982 0.1841
-2.000 0.1996 0.01584 0.00721 -0.0672 0.5900 0.1927
-1.750 0.2260 0.01564 0.00708 -0.0667 0.5793 0.2022
-1.500 0.2531 0.01540 0.00680 -0.0662 0.5709 0.2134
-1.250 0.2797 0.01525 0.00668 -0.0658 0.5599 0.2272
-1.000 0.3068 0.01510 0.00648 -0.0653 0.5516 0.2449
-0.750 0.3328 0.01496 0.00647 -0.0647 0.5407 0.2668
-0.500 0.3595 0.01487 0.00635 -0.0642 0.5328 0.2926
-0.250 0.3857 0.01486 0.00640 -0.0637 0.5228 0.3182
0.000 0.4125 0.01486 0.00632 -0.0632 0.5155 0.3411
0.250 0.4387 0.01488 0.00639 -0.0628 0.5070 0.3608
0.500 0.4652 0.01487 0.00636 -0.0623 0.4996 0.3792
0.750 0.4915 0.01491 0.00639 -0.0619 0.4925 0.3973
1.000 0.5176 0.01492 0.00644 -0.0614 0.4853 0.4157
1.250 0.5443 0.01496 0.00643 -0.0610 0.4795 0.4349
1.500 0.5700 0.01500 0.00655 -0.0605 0.4734 0.4558
1.750 0.5957 0.01500 0.00663 -0.0600 0.4676 0.4817
2.000 0.6207 0.01493 0.00667 -0.0594 0.4626 0.5196
2.250 0.7324 0.01413 0.00699 -0.0757 0.4533 1.0000
2.500 0.7571 0.01436 0.00711 -0.0750 0.4482 1.0000
2.750 0.7823 0.01466 0.00723 -0.0744 0.4439 1.0000
3.000 0.8066 0.01499 0.00749 -0.0737 0.4397 1.0000
3.250 0.8303 0.01526 0.00775 -0.0729 0.4353 1.0000
3.500 0.8545 0.01554 0.00797 -0.0722 0.4311 1.0000
3.750 0.8792 0.01585 0.00817 -0.0715 0.4272 1.0000
4.000 0.9046 0.01628 0.00847 -0.0710 0.4233 1.0000
4.250 0.9268 0.01657 0.00881 -0.0701 0.4194 1.0000
4.500 0.9502 0.01689 0.00914 -0.0693 0.4155 1.0000
4.750 0.9743 0.01722 0.00942 -0.0686 0.4120 1.0000
5.000 0.9993 0.01757 0.00969 -0.0681 0.4088 1.0000
5.250 1.0256 0.01812 0.01012 -0.0678 0.4055 1.0000
5.500 1.0462 0.01845 0.01055 -0.0667 0.4022 1.0000
5.750 1.0680 0.01882 0.01096 -0.0657 0.3986 1.0000
6.000 1.0911 0.01918 0.01132 -0.0649 0.3951 1.0000
6.250 1.1154 0.01955 0.01165 -0.0643 0.3921 1.0000
6.500 1.1412 0.01998 0.01201 -0.0640 0.3894 1.0000
6.750 1.1656 0.02059 0.01260 -0.0636 0.3865 1.0000
7.000 1.1844 0.02102 0.01315 -0.0623 0.3835 1.0000
7.250 1.2046 0.02146 0.01365 -0.0612 0.3800 1.0000
7.500 1.2267 0.02187 0.01410 -0.0603 0.3769 1.0000
7.750 1.2506 0.02229 0.01450 -0.0598 0.3743 1.0000
8.000 1.2764 0.02273 0.01489 -0.0596 0.3718 1.0000
8.250 1.3027 0.02343 0.01554 -0.0595 0.3690 1.0000
8.500 1.3181 0.02395 0.01622 -0.0578 0.3662 1.0000
8.750 1.3348 0.02452 0.01690 -0.0564 0.3630 1.0000
9.000 1.3537 0.02507 0.01753 -0.0553 0.3602 1.0000
9.250 1.3753 0.02554 0.01803 -0.0545 0.3575 1.0000
9.500 1.4002 0.02594 0.01840 -0.0542 0.3550 1.0000
9.750 1.4288 0.02646 0.01885 -0.0545 0.3523 1.0000
10.000 1.4438 0.02727 0.01978 -0.0530 0.3498 1.0000
10.250 1.4540 0.02806 0.02074 -0.0509 0.3470 1.0000
10.500 1.4661 0.02878 0.02157 -0.0491 0.3439 1.0000
10.750 1.4820 0.02936 0.02222 -0.0477 0.3411 1.0000
11.000 1.5044 0.02972 0.02258 -0.0471 0.3384 1.0000
11.250 1.5362 0.02993 0.02270 -0.0477 0.3356 1.0000
11.500 1.5462 0.03073 0.02361 -0.0457 0.3322 1.0000
11.750 1.5412 0.03166 0.02473 -0.0418 0.3286 1.0000
12.000 1.5414 0.03229 0.02543 -0.0384 0.3251 1.0000
12.250 1.5590 0.03245 0.02559 -0.0372 0.3218 1.0000
12.500 1.5992 0.03226 0.02529 -0.0387 0.3186 1.0000
12.750 1.5980 0.03325 0.02639 -0.0355 0.3155 1.0000
13.000 1.5739 0.03509 0.02844 -0.0308 0.3122 1.0000
13.250 1.5657 0.03657 0.03004 -0.0282 0.3086 1.0000
13.500 1.5826 0.03686 0.03032 -0.0273 0.3052 1.0000
13.750 1.6304 0.03573 0.02904 -0.0285 0.3016 1.0000
14.000 1.6087 0.03815 0.03166 -0.0253 0.2984 1.0000
14.250 1.5758 0.04165 0.03539 -0.0225 0.2946 1.0000
14.500 1.5698 0.04372 0.03755 -0.0212 0.2910 1.0000
14.750 1.5976 0.04334 0.03714 -0.0209 0.2879 1.0000
15.000 1.6492 0.04158 0.03524 -0.0215 0.2846 1.0000
15.250 1.5780 0.04860 0.04262 -0.0188 0.2807 1.0000
15.500 1.5256 0.05538 0.04963 -0.0180 0.2759 1.0000
15.750 1.5618 0.05391 0.04813 -0.0175 0.2727 1.0000
16.000 1.6213 0.05049 0.04456 -0.0173 0.2697 1.0000
16.250 1.3550 0.08356 0.07824 -0.0223 0.2560 1.0000
16.500 1.4540 0.07306 0.06768 -0.0191 0.2571 1.0000
16.750 1.5451 0.06458 0.05908 -0.0171 0.2558 1.0000
17.000 1.3344 0.09541 0.09027 -0.0251 0.2400 1.0000
17.250 1.4079 0.08762 0.08247 -0.0223 0.2406 1.0000
17.500 1.5029 0.07744 0.07219 -0.0191 0.2402 1.0000
17.750 1.3216 0.10655 0.10156 -0.0283 0.2232 1.0000
18.000 1.3824 0.10030 0.09531 -0.0259 0.2230 1.0000
18.250 1.5123 0.08452 0.07937 -0.0205 0.2242 1.0000
18.500 1.3917 0.10503 0.10013 -0.0274 0.2114 1.0000
18.750 1.2380 0.13323 0.12846 -0.0387 0.1866 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 702 AIRFOIL (goe702-il)