GOE 767 AIRFOIL (goe767-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: GOE 767 AIRFOIL (goe767-il) Reynolds number: 100,000 Max Cl/Cd: 30.25 at α=6.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe767-il-100000.txt Download as CSV file: xf-goe767-il-100000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 767 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
------ -------- --------- --------- -------- -------- --------
-9.500 -0.7362 0.06847 0.06271 -0.0193 1.0000 0.1019
-9.250 -0.7373 0.06392 0.05809 -0.0193 1.0000 0.1014
-9.000 -0.7469 0.05867 0.05266 -0.0193 1.0000 0.1009
-8.750 -0.7560 0.05331 0.04699 -0.0188 1.0000 0.1007
-8.500 -0.7619 0.04810 0.04135 -0.0177 1.0000 0.1008
-8.250 -0.7620 0.04347 0.03619 -0.0162 1.0000 0.1014
-8.000 -0.7547 0.03964 0.03194 -0.0147 1.0000 0.1026
-7.750 -0.7334 0.03800 0.03041 -0.0140 1.0000 0.1048
-7.500 -0.7145 0.03641 0.02873 -0.0129 1.0000 0.1080
-7.250 -0.7009 0.03383 0.02567 -0.0112 1.0000 0.1123
-7.000 -0.6835 0.03162 0.02323 -0.0099 1.0000 0.1163
-6.750 -0.6612 0.03042 0.02205 -0.0089 1.0000 0.1212
-6.500 -0.6432 0.02850 0.01969 -0.0073 1.0000 0.1283
-6.250 -0.6201 0.02763 0.01899 -0.0064 1.0000 0.1353
-6.000 -0.5995 0.02620 0.01732 -0.0051 1.0000 0.1441
-5.750 -0.5768 0.02544 0.01658 -0.0040 1.0000 0.1529
-5.500 -0.5547 0.02436 0.01544 -0.0029 1.0000 0.1614
-5.250 -0.5320 0.02372 0.01471 -0.0017 1.0000 0.1707
-5.000 -0.5094 0.02278 0.01383 -0.0008 1.0000 0.1792
-4.750 -0.4867 0.02230 0.01322 0.0004 1.0000 0.1890
-4.500 -0.4637 0.02135 0.01243 0.0013 1.0000 0.1977
-4.250 -0.4405 0.02073 0.01172 0.0025 1.0000 0.2060
-4.000 -0.4168 0.01979 0.01086 0.0033 1.0000 0.2128
-3.750 -0.3931 0.01911 0.01025 0.0043 1.0000 0.2199
-3.500 -0.3696 0.01847 0.00957 0.0053 1.0000 0.2270
-3.250 -0.3469 0.01777 0.00909 0.0063 1.0000 0.2347
-3.000 -0.3255 0.01731 0.00868 0.0074 1.0000 0.2432
-2.750 -0.2633 0.01650 0.00813 0.0005 0.9606 0.2610
-2.500 -0.1895 0.01556 0.00744 -0.0075 0.8647 0.2902
-2.250 -0.1545 0.01512 0.00695 -0.0076 0.7655 0.3260
-2.000 -0.1324 0.01462 0.00661 -0.0056 0.7083 0.3931
-1.750 -0.1157 0.01376 0.00655 -0.0026 0.6711 0.5692
-1.500 -0.0941 0.01357 0.00683 0.0007 0.6401 0.7380
-1.250 -0.0619 0.01387 0.00715 0.0021 0.6112 0.8309
-1.000 -0.0200 0.01439 0.00748 0.0013 0.5844 0.8878
-0.750 0.0295 0.01496 0.00780 -0.0012 0.5584 0.9262
-0.500 0.0881 0.01547 0.00803 -0.0060 0.5316 0.9545
-0.250 0.1546 0.01586 0.00809 -0.0128 0.5071 0.9795
0.000 0.2202 0.01601 0.00799 -0.0201 0.4847 1.0000
0.250 0.2397 0.01608 0.00794 -0.0189 0.4710 1.0000
0.500 0.2597 0.01620 0.00798 -0.0178 0.4581 1.0000
0.750 0.2799 0.01641 0.00800 -0.0165 0.4474 1.0000
1.000 0.3007 0.01660 0.00816 -0.0154 0.4357 1.0000
1.250 0.3219 0.01691 0.00834 -0.0142 0.4267 1.0000
1.500 0.3434 0.01717 0.00857 -0.0131 0.4166 1.0000
1.750 0.3652 0.01752 0.00879 -0.0120 0.4086 1.0000
2.000 0.3871 0.01784 0.00912 -0.0109 0.3992 1.0000
2.500 0.4315 0.01855 0.00974 -0.0086 0.3822 1.0000
2.750 0.4542 0.01887 0.00987 -0.0075 0.3744 1.0000
3.000 0.4764 0.01925 0.01034 -0.0064 0.3650 1.0000
3.250 0.4995 0.01957 0.01054 -0.0053 0.3575 1.0000
3.500 0.5220 0.02006 0.01108 -0.0043 0.3494 1.0000
3.750 0.5451 0.02041 0.01140 -0.0032 0.3415 1.0000
4.000 0.5682 0.02091 0.01185 -0.0022 0.3345 1.0000
4.250 0.5909 0.02136 0.01237 -0.0012 0.3260 1.0000
4.500 0.6150 0.02178 0.01259 -0.0002 0.3195 1.0000
4.750 0.6366 0.02235 0.01338 0.0009 0.3102 1.0000
5.000 0.6607 0.02269 0.01357 0.0019 0.3030 1.0000
5.250 0.6822 0.02338 0.01442 0.0030 0.2939 1.0000
5.500 0.7062 0.02374 0.01468 0.0039 0.2864 1.0000
5.750 0.7276 0.02452 0.01560 0.0050 0.2775 1.0000
6.000 0.7515 0.02489 0.01588 0.0060 0.2697 1.0000
6.250 0.7725 0.02579 0.01692 0.0070 0.2612 1.0000
6.500 0.7959 0.02631 0.01741 0.0080 0.2538 1.0000
6.750 0.8171 0.02728 0.01845 0.0090 0.2466 1.0000
7.000 0.8388 0.02804 0.01930 0.0100 0.2390 1.0000
7.250 0.8640 0.02869 0.01975 0.0106 0.2338 1.0000
7.500 0.8795 0.03023 0.02171 0.0119 0.2264 1.0000
7.750 0.9031 0.03083 0.02227 0.0128 0.2208 1.0000
8.000 0.9231 0.03209 0.02360 0.0137 0.2157 1.0000
8.250 0.9375 0.03382 0.02565 0.0149 0.2096 1.0000
8.500 0.9613 0.03439 0.02617 0.0157 0.2048 1.0000
8.750 0.9799 0.03578 0.02763 0.0166 0.2002 1.0000
9.000 0.9872 0.03819 0.03045 0.0180 0.1946 1.0000
9.250 1.0123 0.03841 0.03060 0.0187 0.1897 1.0000
9.500 1.0311 0.03957 0.03177 0.0196 0.1851 1.0000
9.750 1.0285 0.04269 0.03536 0.0212 0.1797 1.0000
10.000 1.0601 0.04193 0.03442 0.0218 0.1745 1.0000
10.250 1.0718 0.04359 0.03619 0.0229 0.1702 1.0000
10.500 1.0498 0.04836 0.04146 0.0246 0.1662 1.0000
10.750 1.0019 0.05522 0.04869 0.0256 0.1643 1.0000
11.000 0.9856 0.05868 0.05222 0.0262 0.1617 1.0000
11.250 0.9562 0.06473 0.05834 0.0238 0.1598 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 767 AIRFOIL (goe767-il)