GOE 604 AIRFOIL (goe604-il) Xfoil prediction polar at RE=100,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: GOE 604 AIRFOIL (goe604-il) Reynolds number: 100,000 Max Cl/Cd: 42.32 at α=5.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe604-il-100000-n5.txt Download as CSV file: xf-goe604-il-100000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 604 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 = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-11.000 -0.4703 0.05117 0.04574 -0.0957 0.9582 0.0513
-10.750 -0.4888 0.04578 0.03987 -0.0990 0.9367 0.0519
-10.500 -0.4953 0.04189 0.03544 -0.0994 0.9155 0.0529
-10.250 -0.4838 0.03935 0.03260 -0.0997 0.8981 0.0538
-10.000 -0.4623 0.03769 0.03085 -0.1001 0.8813 0.0549
-9.750 -0.4438 0.03602 0.02898 -0.0999 0.8632 0.0562
-9.500 -0.4269 0.03435 0.02702 -0.0993 0.8441 0.0577
-9.250 -0.4109 0.03265 0.02492 -0.0983 0.8249 0.0594
-9.000 -0.3916 0.03115 0.02313 -0.0975 0.8063 0.0607
-8.750 -0.3691 0.03004 0.02192 -0.0969 0.7884 0.0620
-8.500 -0.3474 0.02903 0.02073 -0.0962 0.7714 0.0636
-8.250 -0.3260 0.02805 0.01950 -0.0953 0.7557 0.0656
-8.000 -0.3044 0.02708 0.01820 -0.0943 0.7409 0.0678
-7.750 -0.2827 0.02623 0.01731 -0.0934 0.7268 0.0695
-7.500 -0.2606 0.02550 0.01647 -0.0925 0.7143 0.0715
-7.250 -0.2378 0.02480 0.01558 -0.0917 0.7033 0.0742
-7.000 -0.2160 0.02414 0.01477 -0.0906 0.6917 0.0771
-6.750 -0.1938 0.02355 0.01411 -0.0897 0.6819 0.0801
-6.500 -0.1717 0.02300 0.01346 -0.0886 0.6718 0.0838
-6.250 -0.1489 0.02247 0.01281 -0.0877 0.6634 0.0876
-6.000 -0.1268 0.02202 0.01234 -0.0867 0.6540 0.0930
-5.750 -0.1033 0.02160 0.01184 -0.0858 0.6465 0.0998
-5.500 -0.0806 0.02124 0.01147 -0.0849 0.6376 0.1086
-5.250 -0.0565 0.02095 0.01111 -0.0841 0.6300 0.1210
-5.000 -0.0328 0.02071 0.01086 -0.0833 0.6221 0.1365
-4.750 -0.0087 0.02051 0.01060 -0.0825 0.6143 0.1536
-4.500 0.0161 0.02034 0.01032 -0.0818 0.6076 0.1702
-4.250 0.0399 0.02019 0.01015 -0.0810 0.5997 0.1855
-4.000 0.0649 0.02008 0.00993 -0.0803 0.5930 0.2005
-3.750 0.0893 0.01996 0.00980 -0.0796 0.5863 0.2144
-3.500 0.1135 0.01986 0.00969 -0.0788 0.5794 0.2281
-3.250 0.1388 0.01977 0.00954 -0.0782 0.5734 0.2430
-3.000 0.1630 0.01970 0.00949 -0.0775 0.5670 0.2609
-2.750 0.1872 0.01963 0.00943 -0.0767 0.5604 0.2792
-2.500 0.2126 0.01952 0.00927 -0.0761 0.5548 0.2961
-2.250 0.2371 0.01944 0.00919 -0.0754 0.5487 0.3119
-2.000 0.2615 0.01935 0.00911 -0.0746 0.5423 0.3293
-1.750 0.2867 0.01924 0.00899 -0.0740 0.5370 0.3496
-1.500 0.3110 0.01915 0.00895 -0.0732 0.5314 0.3741
-1.250 0.3345 0.01907 0.00897 -0.0724 0.5249 0.4035
-1.000 0.3592 0.01895 0.00890 -0.0716 0.5194 0.4387
-0.750 0.3840 0.01884 0.00889 -0.0709 0.5143 0.4826
-0.500 0.4069 0.01872 0.00901 -0.0699 0.5080 0.5407
-0.250 0.4305 0.01853 0.00908 -0.0688 0.5027 0.6244
0.000 0.4598 0.01833 0.00918 -0.0683 0.4983 0.7401
0.250 0.5099 0.01844 0.00953 -0.0721 0.4909 0.8536
0.500 0.5619 0.01863 0.00963 -0.0764 0.4843 0.9157
0.750 0.6097 0.01884 0.00967 -0.0802 0.4788 0.9530
1.000 0.6611 0.01911 0.00988 -0.0850 0.4716 0.9854
1.250 0.7039 0.01930 0.00992 -0.0882 0.4662 1.0000
1.500 0.7244 0.01948 0.00993 -0.0869 0.4625 1.0000
1.750 0.7409 0.01977 0.01022 -0.0851 0.4572 1.0000
2.000 0.7594 0.02001 0.01040 -0.0835 0.4524 1.0000
2.250 0.7795 0.02022 0.01049 -0.0822 0.4482 1.0000
2.500 0.8002 0.02045 0.01060 -0.0809 0.4445 1.0000
2.750 0.8169 0.02079 0.01095 -0.0791 0.4392 1.0000
3.000 0.8358 0.02107 0.01119 -0.0776 0.4348 1.0000
3.250 0.8567 0.02132 0.01133 -0.0764 0.4311 1.0000
3.500 0.8787 0.02158 0.01149 -0.0754 0.4280 1.0000
3.750 0.8947 0.02201 0.01197 -0.0735 0.4236 1.0000
4.000 0.9130 0.02239 0.01233 -0.0720 0.4196 1.0000
4.250 0.9332 0.02271 0.01261 -0.0708 0.4163 1.0000
4.500 0.9553 0.02301 0.01282 -0.0698 0.4134 1.0000
4.750 0.9757 0.02338 0.01315 -0.0687 0.4103 1.0000
5.000 0.9900 0.02388 0.01372 -0.0666 0.4061 1.0000
5.250 1.0073 0.02428 0.01412 -0.0650 0.4020 1.0000
5.500 1.0277 0.02459 0.01438 -0.0639 0.3986 1.0000
5.750 1.0512 0.02484 0.01454 -0.0632 0.3955 1.0000
6.000 1.0639 0.02540 0.01517 -0.0610 0.3916 1.0000
6.250 1.0772 0.02593 0.01576 -0.0589 0.3875 1.0000
6.500 1.0944 0.02634 0.01616 -0.0574 0.3839 1.0000
6.750 1.1154 0.02662 0.01639 -0.0564 0.3806 1.0000
7.000 1.1358 0.02698 0.01671 -0.0554 0.3776 1.0000
7.250 1.1407 0.02774 0.01759 -0.0522 0.3734 1.0000
7.500 1.1497 0.02832 0.01823 -0.0495 0.3696 1.0000
7.750 1.1643 0.02878 0.01870 -0.0477 0.3665 1.0000
8.000 1.1842 0.02915 0.01903 -0.0467 0.3638 1.0000
8.250 1.2034 0.02958 0.01944 -0.0457 0.3611 1.0000
8.500 1.1984 0.03071 0.02073 -0.0416 0.3574 1.0000
8.750 1.2011 0.03166 0.02177 -0.0387 0.3535 1.0000
9.000 1.2131 0.03227 0.02239 -0.0370 0.3497 1.0000
9.250 1.2350 0.03251 0.02259 -0.0364 0.3465 1.0000
9.500 1.2403 0.03352 0.02366 -0.0342 0.3431 1.0000
9.750 1.2284 0.03543 0.02574 -0.0308 0.3390 1.0000
10.000 1.2286 0.03689 0.02730 -0.0287 0.3352 1.0000
10.250 1.2423 0.03764 0.02806 -0.0277 0.3319 1.0000
10.500 1.2677 0.03773 0.02811 -0.0274 0.3292 1.0000
10.750 1.2453 0.04085 0.03143 -0.0246 0.3250 1.0000
11.000 1.2105 0.04546 0.03624 -0.0222 0.3197 1.0000
11.250 1.2174 0.04694 0.03777 -0.0214 0.3159 1.0000
11.500 1.2470 0.04649 0.03728 -0.0212 0.3133 1.0000
11.750 1.1147 0.06240 0.05358 -0.0206 0.3002 1.0000
12.000 1.1461 0.06132 0.05250 -0.0199 0.2983 1.0000
12.250 1.1853 0.05943 0.05058 -0.0193 0.2968 1.0000
12.750 1.1205 0.07242 0.06379 -0.0203 0.2804 1.0000
13.250 1.0899 0.08211 0.07362 -0.0215 0.2647 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 604 AIRFOIL (goe604-il)