RAE 2822 AIRFOIL (rae2822-il) Xfoil prediction polar at RE=50,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: RAE 2822 AIRFOIL (rae2822-il) Reynolds number: 50,000 Max Cl/Cd: 25.44 at α=5.5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-rae2822-il-50000.txt Download as CSV file: xf-rae2822-il-50000.csv |
XFOIL Version 6.96
Calculated polar for: RAE 2822 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 = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.500 -0.4902 0.11830 0.11070 -0.0136 1.0000 0.3440
-10.250 -0.6088 0.09011 0.08292 -0.0438 1.0000 0.1519
-10.000 -0.6758 0.08073 0.07368 -0.0508 1.0000 0.1402
-9.750 -0.6833 0.07609 0.06902 -0.0507 1.0000 0.1365
-9.500 -0.7035 0.07176 0.06471 -0.0499 1.0000 0.1331
-9.250 -0.7651 0.06740 0.05995 -0.0483 1.0000 0.1245
-9.000 -0.7670 0.06326 0.05570 -0.0469 1.0000 0.1235
-8.750 -0.7710 0.05937 0.05161 -0.0452 1.0000 0.1224
-8.500 -0.7735 0.05557 0.04750 -0.0433 1.0000 0.1213
-8.250 -0.7729 0.05186 0.04346 -0.0414 1.0000 0.1199
-8.000 -0.7687 0.04831 0.03950 -0.0395 1.0000 0.1189
-7.750 -0.7600 0.04502 0.03579 -0.0376 1.0000 0.1191
-7.500 -0.7493 0.04222 0.03251 -0.0358 1.0000 0.1224
-7.250 -0.7354 0.03948 0.02931 -0.0341 1.0000 0.1262
-7.000 -0.7155 0.03687 0.02666 -0.0328 1.0000 0.1312
-6.750 -0.6961 0.03472 0.02410 -0.0311 1.0000 0.1381
-6.500 -0.6751 0.03265 0.02206 -0.0297 1.0000 0.1496
-6.250 -0.6521 0.03076 0.02020 -0.0279 1.0000 0.1621
-6.000 -0.6324 0.02918 0.01863 -0.0258 1.0000 0.1828
-5.750 -0.6161 0.02743 0.01718 -0.0233 1.0000 0.2106
-5.500 -0.6070 0.02516 0.01554 -0.0207 1.0000 0.2590
-5.250 -0.6117 0.02395 0.01667 -0.0132 1.0000 0.5060
-5.000 -0.6084 0.02851 0.02117 -0.0003 1.0000 0.6138
-4.750 -0.5985 0.03140 0.02389 0.0096 1.0000 0.6568
-4.500 -0.5866 0.03325 0.02554 0.0176 1.0000 0.6928
-4.250 -0.5702 0.03471 0.02680 0.0246 1.0000 0.7261
-4.000 -0.5341 0.03647 0.02824 0.0298 1.0000 0.7657
-3.750 -0.4217 0.03811 0.02921 0.0229 1.0000 0.8242
-3.250 -0.3497 0.03655 0.02715 0.0192 1.0000 0.8612
-3.000 -0.3504 0.03577 0.02632 0.0227 1.0000 0.8708
-2.750 -0.3240 0.03489 0.02526 0.0215 1.0000 0.8793
-2.500 -0.3393 0.03412 0.02452 0.0272 1.0000 0.8858
-2.250 -0.3144 0.03331 0.02356 0.0261 1.0000 0.8920
-2.000 -0.3184 0.03254 0.02277 0.0298 1.0000 0.8982
-1.750 -0.3100 0.03179 0.02195 0.0314 1.0000 0.9039
-1.500 -0.2989 0.03110 0.02120 0.0326 1.0000 0.9101
-1.250 -0.3005 0.03035 0.02043 0.0359 1.0000 0.9164
-1.000 -0.2743 0.02982 0.01980 0.0343 1.0000 0.9224
-0.750 -0.2737 0.02915 0.01910 0.0372 1.0000 0.9295
-0.500 -0.2386 0.02879 0.01867 0.0340 1.0000 0.9353
-0.250 -0.2271 0.02831 0.01816 0.0349 1.0000 0.9432
0.000 -0.1879 0.02810 0.01789 0.0308 1.0000 0.9493
0.250 -0.1661 0.02782 0.01759 0.0298 1.0000 0.9573
0.500 -0.1256 0.02773 0.01749 0.0254 1.0000 0.9643
0.750 -0.0939 0.02764 0.01741 0.0224 1.0000 0.9724
1.000 -0.0565 0.02764 0.01743 0.0183 1.0000 0.9810
1.250 -0.0187 0.02769 0.01753 0.0140 1.0000 0.9902
1.500 0.0134 0.02780 0.01769 0.0106 1.0000 1.0000
1.750 0.0088 0.02745 0.01738 0.0136 1.0000 1.0000
2.000 0.0031 0.02702 0.01700 0.0167 1.0000 1.0000
2.250 -0.0029 0.02653 0.01656 0.0198 1.0000 1.0000
2.500 -0.0056 0.02609 0.01615 0.0222 1.0000 1.0000
2.750 0.0038 0.02601 0.01610 0.0225 1.0000 1.0000
3.000 0.0220 0.02625 0.01639 0.0213 1.0000 1.0000
3.250 0.0443 0.02671 0.01692 0.0194 1.0000 1.0000
3.500 0.0682 0.02734 0.01761 0.0172 1.0000 1.0000
3.750 0.0923 0.02809 0.01844 0.0150 1.0000 1.0000
4.000 0.1233 0.02921 0.01967 0.0115 0.9962 1.0000
4.250 0.1959 0.03172 0.02245 0.0007 0.9700 1.0000
4.500 0.2596 0.03346 0.02448 -0.0078 0.9414 1.0000
4.750 0.3174 0.03459 0.02595 -0.0143 0.9088 1.0000
5.000 0.3957 0.03488 0.02673 -0.0217 0.8616 1.0000
5.250 0.5766 0.02598 0.01929 -0.0281 0.7430 1.0000
5.500 0.6025 0.02368 0.01436 -0.0147 0.2838 1.0000
5.750 0.6231 0.02637 0.01620 -0.0139 0.2097 1.0000
6.000 0.6692 0.02884 0.01836 -0.0163 0.1700 1.0000
6.250 0.7147 0.03137 0.02075 -0.0189 0.1486 1.0000
6.500 0.7522 0.03399 0.02360 -0.0200 0.1378 1.0000
6.750 0.7835 0.03686 0.02662 -0.0206 0.1299 1.0000
7.000 0.8083 0.03948 0.02967 -0.0201 0.1238 1.0000
7.250 0.8321 0.04272 0.03334 -0.0194 0.1222 1.0000
7.500 0.8520 0.04610 0.03707 -0.0185 0.1208 1.0000
7.750 0.8720 0.05023 0.04129 -0.0182 0.1179 1.0000
8.000 0.8832 0.05391 0.04545 -0.0164 0.1178 1.0000
8.250 0.8779 0.05760 0.05009 -0.0129 0.1234 1.0000
8.500 0.8785 0.06243 0.05533 -0.0112 0.1286 1.0000
8.750 0.8719 0.06734 0.06069 -0.0094 0.1364 1.0000
9.000 0.8773 0.07355 0.06710 -0.0092 0.1476 1.0000
9.250 0.8267 0.07856 0.07250 -0.0076 0.1563 1.0000
9.500 0.8113 0.08804 0.08213 -0.0104 0.1896 1.0000
|
Polar data table (+)
Polar graphs
<< Back to RAE 2822 AIRFOIL (rae2822-il)