GOE 804 (EA 8) AIRFOIL (goe804-il) Xfoil prediction polar at RE=500,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: GOE 804 (EA 8) AIRFOIL (goe804-il) Reynolds number: 500,000 Max Cl/Cd: 169.11 at α=3° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe804-il-500000.txt Download as CSV file: xf-goe804-il-500000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 804 (EA 8) AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.500 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.000 -0.2099 0.09449 0.09212 -0.0736 0.9672 0.0165
-7.750 -0.2008 0.09135 0.08899 -0.0763 0.9643 0.0165
-7.500 -0.1922 0.08650 0.08415 -0.0777 0.9629 0.0170
-7.250 -0.1767 0.08393 0.08158 -0.0787 0.9620 0.0177
-7.000 -0.1610 0.08101 0.07865 -0.0821 0.9607 0.0183
-6.750 -0.1707 0.07982 0.07749 -0.0786 0.9530 0.0187
-6.500 -0.1525 0.07628 0.07394 -0.0831 0.9505 0.0196
-6.250 -0.1264 0.07202 0.06967 -0.0909 0.9488 0.0214
-6.000 -0.1143 0.06916 0.06681 -0.0971 0.9413 0.0223
-5.750 -0.0866 0.06488 0.06249 -0.1046 0.9386 0.0224
-5.500 -0.0619 0.05801 0.05560 -0.1129 0.9369 0.0231
-5.250 -0.0350 0.05498 0.05253 -0.1167 0.9360 0.0238
-5.000 0.0003 0.05140 0.04891 -0.1232 0.9351 0.0248
-4.750 0.0133 0.04882 0.04629 -0.1244 0.9276 0.0258
-4.500 0.0812 0.04242 0.03963 -0.1394 0.9263 0.0300
-4.250 0.2160 0.01702 0.01227 -0.1764 0.9357 0.0235
-4.000 0.2553 0.01609 0.01118 -0.1792 0.9353 0.0254
-3.750 0.2941 0.01505 0.00990 -0.1818 0.9350 0.0278
-3.500 0.3256 0.01452 0.00918 -0.1827 0.9336 0.0306
-3.250 0.3440 0.01403 0.00862 -0.1811 0.9287 0.0333
-3.000 0.3767 0.01375 0.00828 -0.1821 0.9272 0.0370
-2.750 0.4128 0.01314 0.00756 -0.1840 0.9263 0.0420
-2.500 0.4498 0.01273 0.00715 -0.1859 0.9252 0.0476
-2.250 0.4882 0.01247 0.00683 -0.1879 0.9243 0.0520
-2.000 0.5310 0.01152 0.00586 -0.1913 0.9240 0.0603
-1.750 0.5714 0.01113 0.00543 -0.1938 0.9234 0.0652
-1.500 0.6103 0.01061 0.00494 -0.1963 0.9229 0.0710
-1.250 0.6474 0.01033 0.00466 -0.1983 0.9222 0.0762
-1.000 0.6868 0.00992 0.00425 -0.2008 0.9213 0.0811
-0.750 0.7063 0.00989 0.00427 -0.1989 0.9151 0.0871
-0.500 0.7413 0.00960 0.00402 -0.2004 0.9129 0.0992
-0.250 0.7776 0.00925 0.00376 -0.2022 0.9111 0.1267
0.000 0.8172 0.00878 0.00353 -0.2049 0.9092 0.2182
0.250 0.8610 0.00749 0.00365 -0.2089 0.9075 0.7934
0.500 0.8772 0.00739 0.00369 -0.2059 0.8995 0.8782
0.750 0.9053 0.00704 0.00337 -0.2053 0.8941 0.9245
1.000 0.9203 0.00686 0.00324 -0.2020 0.8858 1.0000
1.250 0.9540 0.00675 0.00310 -0.2031 0.8806 1.0000
1.500 0.9798 0.00676 0.00312 -0.2026 0.8727 1.0000
1.750 1.0117 0.00668 0.00301 -0.2033 0.8659 1.0000
2.000 1.0363 0.00671 0.00306 -0.2025 0.8562 1.0000
2.250 1.0632 0.00669 0.00303 -0.2021 0.8444 1.0000
2.500 1.0898 0.00668 0.00304 -0.2016 0.8305 1.0000
2.750 1.1160 0.00670 0.00304 -0.2011 0.8133 1.0000
3.000 1.1415 0.00675 0.00300 -0.2003 0.7808 1.0000
3.250 1.1639 0.00703 0.00303 -0.1988 0.7265 1.0000
3.500 1.1807 0.00763 0.00329 -0.1963 0.6612 1.0000
3.750 1.1929 0.00846 0.00371 -0.1929 0.5846 1.0000
4.000 1.1970 0.00978 0.00434 -0.1882 0.4550 1.0000
4.250 1.2100 0.01084 0.00494 -0.1854 0.3743 1.0000
4.500 1.2226 0.01206 0.00557 -0.1828 0.2625 1.0000
4.750 1.2405 0.01291 0.00614 -0.1811 0.2127 1.0000
5.000 1.2592 0.01373 0.00668 -0.1795 0.1495 1.0000
5.250 1.2756 0.01478 0.00739 -0.1776 0.0925 1.0000
5.500 1.2947 0.01553 0.00796 -0.1761 0.0647 1.0000
5.750 1.3149 0.01614 0.00856 -0.1747 0.0529 1.0000
6.000 1.3343 0.01682 0.00923 -0.1732 0.0407 1.0000
6.250 1.3534 0.01754 0.00993 -0.1716 0.0313 1.0000
6.500 1.3711 0.01839 0.01076 -0.1697 0.0254 1.0000
6.750 1.3895 0.01912 0.01157 -0.1679 0.0231 1.0000
7.000 1.4077 0.01988 0.01239 -0.1662 0.0210 1.0000
7.250 1.4236 0.02086 0.01341 -0.1641 0.0189 1.0000
7.500 1.4351 0.02235 0.01504 -0.1611 0.0173 1.0000
7.750 1.4525 0.02320 0.01597 -0.1593 0.0164 1.0000
8.000 1.4686 0.02425 0.01712 -0.1573 0.0155 1.0000
8.250 1.4844 0.02538 0.01835 -0.1553 0.0146 1.0000
8.500 1.5000 0.02656 0.01960 -0.1533 0.0139 1.0000
8.750 1.5150 0.02802 0.02114 -0.1513 0.0132 1.0000
9.000 1.5360 0.03185 0.02522 -0.1506 0.0122 1.0000
9.250 1.5506 0.03258 0.02609 -0.1484 0.0118 1.0000
9.500 1.5671 0.03400 0.02767 -0.1467 0.0113 1.0000
9.750 1.5840 0.03603 0.02992 -0.1451 0.0108 1.0000
10.000 1.5988 0.03829 0.03242 -0.1432 0.0105 1.0000
10.250 1.6105 0.04095 0.03535 -0.1409 0.0103 1.0000
10.500 1.6176 0.04356 0.03822 -0.1380 0.0100 1.0000
10.750 1.6213 0.04613 0.04102 -0.1349 0.0097 1.0000
11.000 1.6212 0.04906 0.04422 -0.1314 0.0096 1.0000
11.250 1.6143 0.05300 0.04850 -0.1273 0.0095 1.0000
11.500 1.5982 0.05808 0.05400 -0.1226 0.0096 1.0000
11.750 1.5701 0.06489 0.06129 -0.1176 0.0098 1.0000
12.000 1.5345 0.07276 0.06959 -0.1134 0.0101 1.0000
12.250 1.5020 0.08024 0.07740 -0.1109 0.0103 1.0000
|
Polar data table (+)
Polar graphs
<< Back to GOE 804 (EA 8) AIRFOIL (goe804-il)