GOE 685 AIRFOIL (goe685-il) Xfoil prediction polar at RE=50,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: GOE 685 AIRFOIL (goe685-il) Reynolds number: 50,000 Max Cl/Cd: 33.89 at α=7.25° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe685-il-50000-n5.txt Download as CSV file: xf-goe685-il-50000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 685 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 -0.3467 0.10210 0.09429 -0.0453 1.0000 0.0896
-9.750 -0.3513 0.09865 0.09091 -0.0458 1.0000 0.0902
-9.500 -0.3582 0.09508 0.08742 -0.0462 1.0000 0.0905
-9.250 -0.3677 0.09143 0.08386 -0.0464 1.0000 0.0909
-9.000 -0.3804 0.08779 0.08033 -0.0464 1.0000 0.0912
-8.750 -0.3979 0.08411 0.07677 -0.0460 1.0000 0.0915
-8.500 -0.4209 0.08058 0.07338 -0.0450 1.0000 0.0917
-8.000 -0.4818 0.07166 0.06475 -0.0443 1.0000 0.0923
-7.750 -0.5299 0.06040 0.05348 -0.0506 1.0000 0.0927
-7.500 -0.5672 0.04717 0.03958 -0.0592 1.0000 0.0947
-7.250 -0.5609 0.04471 0.03698 -0.0591 1.0000 0.0973
-7.000 -0.5413 0.04373 0.03598 -0.0594 0.9973 0.1009
-6.750 -0.5142 0.03836 0.02986 -0.0654 0.9898 0.1078
-6.500 -0.4801 0.03804 0.02964 -0.0673 0.9828 0.1129
-6.250 -0.4496 0.03590 0.02716 -0.0701 0.9749 0.1225
-6.000 -0.4160 0.03453 0.02551 -0.0729 0.9678 0.1348
-5.750 -0.3837 0.03387 0.02471 -0.0747 0.9595 0.1481
-5.500 -0.3508 0.03368 0.02445 -0.0762 0.9515 0.1624
-5.250 -0.3169 0.03362 0.02429 -0.0779 0.9431 0.1784
-5.000 -0.2854 0.03400 0.02465 -0.0787 0.9341 0.1929
-4.750 -0.2499 0.03426 0.02481 -0.0803 0.9257 0.2084
-4.500 -0.2196 0.03446 0.02498 -0.0807 0.9156 0.2205
-4.250 -0.1815 0.03416 0.02452 -0.0828 0.9075 0.2337
-4.000 -0.1522 0.03350 0.02360 -0.0836 0.8962 0.2453
-3.750 -0.1108 0.03300 0.02291 -0.0861 0.8890 0.2567
-3.250 -0.0476 0.03181 0.02134 -0.0879 0.8686 0.2748
-3.000 -0.0167 0.03130 0.02076 -0.0885 0.8584 0.2814
-2.750 0.0144 0.03075 0.02002 -0.0892 0.8486 0.2900
-2.500 0.0489 0.03014 0.01928 -0.0903 0.8400 0.2977
-2.250 0.0763 0.02969 0.01872 -0.0902 0.8288 0.3046
-2.000 0.1141 0.02905 0.01791 -0.0918 0.8215 0.3137
-1.750 0.1396 0.02875 0.01755 -0.0913 0.8094 0.3213
-1.500 0.1723 0.02825 0.01689 -0.0920 0.7999 0.3301
-1.250 0.2038 0.02781 0.01642 -0.0923 0.7900 0.3381
-1.000 0.2317 0.02752 0.01600 -0.0921 0.7787 0.3476
-0.750 0.2681 0.02700 0.01543 -0.0930 0.7709 0.3580
-0.500 0.2947 0.02679 0.01514 -0.0926 0.7584 0.3668
-0.250 0.3261 0.02651 0.01477 -0.0928 0.7480 0.3777
0.000 0.3599 0.02614 0.01438 -0.0933 0.7384 0.3888
0.250 0.3876 0.02601 0.01419 -0.0931 0.7262 0.4001
0.500 0.4214 0.02571 0.01382 -0.0936 0.7167 0.4137
0.750 0.4498 0.02554 0.01369 -0.0934 0.7051 0.4260
1.000 0.4773 0.02545 0.01363 -0.0931 0.6933 0.4395
1.250 0.5111 0.02514 0.01334 -0.0935 0.6844 0.4560
1.750 0.5624 0.02512 0.01349 -0.0924 0.6608 0.4912
2.000 0.5915 0.02497 0.01346 -0.0923 0.6510 0.5156
2.250 0.6144 0.02498 0.01372 -0.0914 0.6396 0.5484
2.500 0.6425 0.02445 0.01363 -0.0906 0.6317 0.6294
2.750 0.6765 0.02428 0.01395 -0.0917 0.6198 1.0000
3.000 0.7047 0.02459 0.01409 -0.0916 0.6110 1.0000
3.250 0.7296 0.02502 0.01442 -0.0910 0.6013 1.0000
3.500 0.7550 0.02544 0.01476 -0.0906 0.5927 1.0000
3.750 0.7802 0.02588 0.01512 -0.0900 0.5840 1.0000
4.000 0.8049 0.02635 0.01555 -0.0894 0.5759 1.0000
4.250 0.8291 0.02680 0.01597 -0.0888 0.5672 1.0000
4.500 0.8529 0.02727 0.01641 -0.0880 0.5586 1.0000
4.750 0.8774 0.02767 0.01680 -0.0873 0.5497 1.0000
5.000 0.8991 0.02819 0.01732 -0.0862 0.5405 1.0000
5.250 0.9252 0.02848 0.01758 -0.0856 0.5317 1.0000
5.500 0.9441 0.02910 0.01826 -0.0842 0.5218 1.0000
5.750 0.9730 0.02923 0.01834 -0.0838 0.5134 1.0000
6.000 0.9886 0.02997 0.01916 -0.0821 0.5027 1.0000
6.250 1.0157 0.03017 0.01935 -0.0815 0.4940 1.0000
6.500 1.0326 0.03087 0.02014 -0.0799 0.4839 1.0000
6.750 1.0566 0.03131 0.02061 -0.0791 0.4759 1.0000
7.000 1.0738 0.03212 0.02153 -0.0776 0.4674 1.0000
7.250 1.1010 0.03249 0.02195 -0.0772 0.4606 1.0000
7.500 1.1127 0.03356 0.02318 -0.0752 0.4515 1.0000
7.750 1.1419 0.03379 0.02343 -0.0749 0.4441 1.0000
8.000 1.1500 0.03493 0.02476 -0.0725 0.4339 1.0000
8.250 1.1721 0.03535 0.02525 -0.0713 0.4246 1.0000
8.500 1.1875 0.03595 0.02595 -0.0694 0.4135 1.0000
8.750 1.1965 0.03685 0.02698 -0.0669 0.4022 1.0000
9.000 1.2149 0.03714 0.02731 -0.0652 0.3906 1.0000
9.250 1.2239 0.03789 0.02818 -0.0626 0.3789 1.0000
9.500 1.2232 0.03906 0.02949 -0.0592 0.3675 1.0000
9.750 1.2323 0.03969 0.03019 -0.0566 0.3560 1.0000
10.000 1.2366 0.04056 0.03113 -0.0538 0.3438 1.0000
10.250 1.2285 0.04245 0.03318 -0.0506 0.3319 1.0000
10.500 1.2262 0.04401 0.03484 -0.0481 0.3191 1.0000
10.750 1.2239 0.04564 0.03657 -0.0458 0.3054 1.0000
11.000 1.2209 0.04753 0.03853 -0.0440 0.2911 1.0000
11.250 1.2178 0.04967 0.04075 -0.0424 0.2769 1.0000
11.500 1.2148 0.05203 0.04318 -0.0412 0.2627 1.0000
11.750 1.2097 0.05480 0.04604 -0.0404 0.2475 1.0000
12.000 1.2031 0.05789 0.04918 -0.0398 0.2307 1.0000
12.250 1.1985 0.06087 0.05215 -0.0393 0.2136 1.0000
12.500 1.1946 0.06383 0.05503 -0.0389 0.1966 1.0000
12.750 1.1895 0.06707 0.05816 -0.0386 0.1806 1.0000
13.000 1.1824 0.07077 0.06180 -0.0387 0.1655 1.0000
13.250 1.1745 0.07481 0.06580 -0.0390 0.1511 1.0000
13.500 1.1672 0.07899 0.07000 -0.0395 0.1377 1.0000
13.750 1.1615 0.08311 0.07416 -0.0402 0.1261 1.0000
14.000 1.1573 0.08707 0.07809 -0.0408 0.1168 1.0000
14.250 1.1545 0.09099 0.08207 -0.0416 0.1090 1.0000
14.500 1.1526 0.09471 0.08574 -0.0423 0.1032 1.0000
14.750 1.1520 0.09827 0.08929 -0.0431 0.0983 1.0000
15.000 1.1535 0.10145 0.09239 -0.0436 0.0946 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 685 AIRFOIL (goe685-il)