GOE 679 AIRFOIL (goe679-il) Xfoil prediction polar at RE=200,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: GOE 679 AIRFOIL (goe679-il) Reynolds number: 200,000 Max Cl/Cd: 58.07 at α=5.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe679-il-200000.txt Download as CSV file: xf-goe679-il-200000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 679 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
------ -------- --------- --------- -------- -------- --------
-12.750 -0.8101 0.05556 0.05050 -0.0858 1.0000 0.0598
-12.500 -0.8448 0.05332 0.04816 -0.0819 1.0000 0.0597
-12.250 -0.8786 0.05184 0.04662 -0.0762 1.0000 0.0596
-12.000 -0.8902 0.04883 0.04328 -0.0761 0.9939 0.0597
-11.750 -0.8856 0.04557 0.03968 -0.0771 0.9859 0.0599
-11.500 -0.8629 0.04307 0.03705 -0.0792 0.9814 0.0604
-11.250 -0.8445 0.04138 0.03530 -0.0794 0.9744 0.0610
-11.000 -0.8198 0.03972 0.03351 -0.0808 0.9696 0.0617
-10.750 -0.8012 0.03819 0.03185 -0.0807 0.9625 0.0625
-10.500 -0.7772 0.03659 0.03007 -0.0816 0.9571 0.0634
-10.250 -0.7560 0.03504 0.02828 -0.0817 0.9515 0.0642
-10.000 -0.7365 0.03356 0.02658 -0.0813 0.9445 0.0649
-9.750 -0.7062 0.03214 0.02486 -0.0826 0.9411 0.0658
-9.500 -0.6884 0.03093 0.02365 -0.0813 0.9333 0.0666
-9.250 -0.6576 0.02988 0.02263 -0.0823 0.9289 0.0677
-9.000 -0.6213 0.02882 0.02153 -0.0842 0.9261 0.0692
-8.750 -0.6041 0.02800 0.02062 -0.0825 0.9171 0.0707
-8.500 -0.5700 0.02698 0.01940 -0.0838 0.9127 0.0725
-8.250 -0.5310 0.02581 0.01836 -0.0861 0.9101 0.0750
-8.000 -0.5141 0.02517 0.01770 -0.0840 0.9003 0.0771
-7.750 -0.4787 0.02412 0.01657 -0.0854 0.8964 0.0802
-7.500 -0.4401 0.02314 0.01565 -0.0873 0.8940 0.0842
-7.250 -0.4254 0.02262 0.01507 -0.0848 0.8844 0.0879
-7.000 -0.3919 0.02180 0.01430 -0.0858 0.8805 0.0939
-6.750 -0.3541 0.02094 0.01347 -0.0874 0.8782 0.1040
-6.500 -0.3275 0.02039 0.01295 -0.0870 0.8725 0.1160
-6.250 -0.3023 0.01999 0.01258 -0.0863 0.8656 0.1284
-6.000 -0.2635 0.01949 0.01209 -0.0880 0.8627 0.1419
-5.750 -0.2204 0.01900 0.01155 -0.0906 0.8605 0.1539
-5.500 -0.1941 0.01883 0.01135 -0.0900 0.8532 0.1625
-5.250 -0.1602 0.01841 0.01095 -0.0909 0.8474 0.1707
-5.000 -0.1133 0.01803 0.01042 -0.0942 0.8440 0.1804
-4.750 -0.0586 0.01743 0.00992 -0.0993 0.8411 0.1901
-4.500 -0.0399 0.01735 0.00971 -0.0972 0.8303 0.1963
-4.250 0.0095 0.01689 0.00933 -0.1013 0.8249 0.2066
-4.000 0.0423 0.01670 0.00908 -0.1020 0.8159 0.2165
-3.750 0.0793 0.01648 0.00888 -0.1035 0.8068 0.2260
-3.500 0.1137 0.01622 0.00851 -0.1046 0.7975 0.2338
-3.250 0.1454 0.01591 0.00823 -0.1051 0.7870 0.2409
-3.000 0.1740 0.01577 0.00800 -0.1049 0.7758 0.2486
-2.750 0.2079 0.01544 0.00765 -0.1058 0.7650 0.2567
-2.500 0.2283 0.01533 0.00753 -0.1040 0.7521 0.2642
-2.250 0.2615 0.01512 0.00723 -0.1047 0.7411 0.2740
-2.000 0.2804 0.01496 0.00710 -0.1026 0.7273 0.2826
-1.750 0.3029 0.01487 0.00692 -0.1012 0.7136 0.2926
-1.500 0.3303 0.01465 0.00671 -0.1008 0.7008 0.3048
-1.250 0.3487 0.01452 0.00656 -0.0986 0.6860 0.3169
-1.000 0.3674 0.01440 0.00646 -0.0964 0.6710 0.3301
-0.750 0.3885 0.01426 0.00632 -0.0948 0.6565 0.3458
-0.500 0.4116 0.01414 0.00616 -0.0936 0.6426 0.3644
-0.250 0.4270 0.01403 0.00610 -0.0908 0.6273 0.3837
0.000 0.4444 0.01392 0.00605 -0.0885 0.6129 0.4080
0.250 0.4638 0.01381 0.00599 -0.0866 0.5999 0.4421
0.750 0.4913 0.01337 0.00604 -0.0805 0.5749 0.5773
1.000 0.5109 0.01300 0.00621 -0.0781 0.5641 0.7510
1.250 0.6258 0.01358 0.00693 -0.0951 0.5452 0.9209
1.500 0.6798 0.01406 0.00731 -0.1000 0.5316 0.9542
1.750 0.7335 0.01453 0.00761 -0.1050 0.5198 0.9749
2.000 0.7936 0.01493 0.00782 -0.1115 0.5080 0.9894
2.250 0.8514 0.01519 0.00798 -0.1178 0.4967 1.0000
2.500 0.8656 0.01536 0.00803 -0.1153 0.4891 1.0000
2.750 0.8799 0.01555 0.00814 -0.1128 0.4820 1.0000
3.000 0.8908 0.01568 0.00824 -0.1095 0.4745 1.0000
3.250 0.9098 0.01593 0.00833 -0.1079 0.4676 1.0000
3.500 0.9210 0.01609 0.00849 -0.1047 0.4608 1.0000
3.750 0.9335 0.01626 0.00863 -0.1017 0.4540 1.0000
4.000 0.9559 0.01656 0.00876 -0.1007 0.4473 1.0000
4.250 0.9644 0.01672 0.00896 -0.0969 0.4408 1.0000
4.500 0.9769 0.01691 0.00911 -0.0939 0.4342 1.0000
4.750 1.0010 0.01724 0.00928 -0.0932 0.4277 1.0000
5.000 1.0073 0.01740 0.00950 -0.0890 0.4216 1.0000
5.250 1.0178 0.01759 0.00967 -0.0855 0.4153 1.0000
5.500 1.0407 0.01792 0.00986 -0.0847 0.4090 1.0000
5.750 1.0464 0.01810 0.01009 -0.0803 0.4028 1.0000
6.000 1.0572 0.01831 0.01029 -0.0770 0.3964 1.0000
6.250 1.0790 0.01865 0.01049 -0.0760 0.3901 1.0000
6.500 1.0873 0.01890 0.01081 -0.0723 0.3840 1.0000
6.750 1.0990 0.01916 0.01106 -0.0693 0.3776 1.0000
7.000 1.1196 0.01951 0.01129 -0.0681 0.3716 1.0000
7.250 1.1300 0.01983 0.01167 -0.0649 0.3655 1.0000
7.500 1.1412 0.02013 0.01198 -0.0620 0.3591 1.0000
7.750 1.1612 0.02050 0.01222 -0.0607 0.3531 1.0000
8.000 1.1710 0.02089 0.01268 -0.0576 0.3470 1.0000
8.250 1.1827 0.02125 0.01306 -0.0550 0.3408 1.0000
8.500 1.2024 0.02166 0.01334 -0.0537 0.3351 1.0000
8.750 1.2133 0.02212 0.01388 -0.0510 0.3294 1.0000
9.000 1.2254 0.02257 0.01435 -0.0486 0.3237 1.0000
9.250 1.2431 0.02301 0.01472 -0.0472 0.3187 1.0000
9.500 1.2587 0.02356 0.01528 -0.0455 0.3135 1.0000
9.750 1.2700 0.02410 0.01588 -0.0431 0.3084 1.0000
10.000 1.2837 0.02459 0.01636 -0.0412 0.3034 1.0000
10.250 1.3063 0.02514 0.01679 -0.0407 0.2986 1.0000
10.500 1.3123 0.02580 0.01758 -0.0377 0.2939 1.0000
10.750 1.3241 0.02641 0.01822 -0.0357 0.2893 1.0000
11.000 1.3405 0.02697 0.01873 -0.0344 0.2851 1.0000
11.250 1.3594 0.02762 0.01935 -0.0335 0.2809 1.0000
11.500 1.3651 0.02843 0.02029 -0.0309 0.2767 1.0000
11.750 1.3764 0.02916 0.02105 -0.0291 0.2726 1.0000
12.000 1.3919 0.02980 0.02167 -0.0278 0.2690 1.0000
12.250 1.4137 0.03044 0.02223 -0.0274 0.2650 1.0000
12.500 1.4164 0.03148 0.02343 -0.0248 0.2616 1.0000
12.750 1.4245 0.03243 0.02446 -0.0230 0.2579 1.0000
13.000 1.4369 0.03324 0.02529 -0.0216 0.2546 1.0000
13.250 1.4561 0.03388 0.02587 -0.0210 0.2514 1.0000
13.500 1.4706 0.03480 0.02682 -0.0200 0.2482 1.0000
13.750 1.4706 0.03618 0.02837 -0.0177 0.2451 1.0000
14.000 1.4766 0.03738 0.02965 -0.0160 0.2420 1.0000
14.250 1.4873 0.03838 0.03068 -0.0149 0.2390 1.0000
14.500 1.5036 0.03916 0.03144 -0.0141 0.2363 1.0000
14.750 1.5256 0.03987 0.03211 -0.0139 0.2332 1.0000
15.000 1.5186 0.04182 0.03427 -0.0117 0.2308 1.0000
15.250 1.5184 0.04357 0.03617 -0.0101 0.2281 1.0000
15.500 1.5225 0.04509 0.03776 -0.0088 0.2252 1.0000
15.750 1.5333 0.04617 0.03885 -0.0080 0.2222 1.0000
16.000 1.5579 0.04654 0.03913 -0.0079 0.2193 1.0000
16.250 1.5530 0.04879 0.04154 -0.0065 0.2168 1.0000
16.500 1.5406 0.05164 0.04460 -0.0050 0.2142 1.0000
16.750 1.5359 0.05407 0.04716 -0.0040 0.2114 1.0000
17.000 1.5403 0.05580 0.04894 -0.0034 0.2085 1.0000
17.250 1.5573 0.05650 0.04959 -0.0030 0.2057 1.0000
17.500 1.5698 0.05773 0.05083 -0.0025 0.2028 1.0000
17.750 1.5410 0.06252 0.05590 -0.0017 0.2003 1.0000
18.000 1.5225 0.06672 0.06028 -0.0015 0.1973 1.0000
18.250 1.5212 0.06925 0.06288 -0.0014 0.1942 1.0000
18.500 1.5420 0.06944 0.06301 -0.0011 0.1913 1.0000
18.750 1.5465 0.07147 0.06506 -0.0009 0.1883 1.0000
19.000 1.4930 0.08012 0.07407 -0.0021 0.1851 1.0000
19.250 1.4666 0.08609 0.08020 -0.0033 0.1814 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 679 AIRFOIL (goe679-il)