RUTAN CANARD AIRFOIL (amsoil1-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: RUTAN CANARD AIRFOIL (amsoil1-il) Reynolds number: 100,000 Max Cl/Cd: 48.61 at α=5° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-amsoil1-il-100000.txt Download as CSV file: xf-amsoil1-il-100000.csv |
XFOIL Version 6.96
Calculated polar for: RUTAN CANARD 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 -0.4492 0.10596 0.10062 -0.0231 1.0000 0.1257
-9.250 -0.4843 0.10217 0.09696 -0.0335 1.0000 0.1302
-9.000 -0.4762 0.09701 0.09185 -0.0334 1.0000 0.1321
-8.750 -0.4497 0.09424 0.08905 -0.0297 1.0000 0.1351
-8.500 -0.4425 0.09112 0.08596 -0.0303 1.0000 0.1392
-8.250 -0.4624 0.08702 0.08197 -0.0362 1.0000 0.1447
-8.000 -0.5294 0.08484 0.07979 -0.0406 1.0000 0.1466
-7.750 -0.5441 0.08117 0.07619 -0.0380 1.0000 0.1478
-7.500 -0.5138 0.07783 0.07300 -0.0343 1.0000 0.1505
-7.250 -0.5180 0.07615 0.07138 -0.0307 1.0000 0.1529
-7.000 -0.5311 0.07434 0.06958 -0.0279 1.0000 0.1561
-6.750 -0.5819 0.07396 0.06875 -0.0284 1.0000 0.1641
-6.500 -0.5738 0.06864 0.06366 -0.0263 1.0000 0.1664
-6.250 -0.5643 0.06641 0.06156 -0.0233 1.0000 0.1699
-6.000 -0.5674 0.06463 0.05961 -0.0227 1.0000 0.1799
-5.750 -0.5670 0.06099 0.05593 -0.0217 1.0000 0.1858
-5.500 -0.5585 0.05878 0.05376 -0.0198 1.0000 0.1909
-5.250 -0.5358 0.05516 0.04995 -0.0231 0.9958 0.2059
-5.000 -0.4924 0.04098 0.03379 -0.0318 0.9916 0.1180
-4.750 -0.4591 0.03737 0.02974 -0.0338 0.9875 0.1111
-4.500 -0.4231 0.03431 0.02619 -0.0361 0.9841 0.1094
-4.250 -0.3903 0.03229 0.02377 -0.0376 0.9795 0.1112
-4.000 -0.3547 0.03041 0.02147 -0.0394 0.9752 0.1124
-3.750 -0.3149 0.02917 0.01974 -0.0417 0.9716 0.1154
-3.500 -0.2853 0.02767 0.01825 -0.0427 0.9670 0.1197
-3.250 -0.2505 0.02674 0.01722 -0.0443 0.9623 0.1244
-3.000 -0.2104 0.02595 0.01619 -0.0467 0.9586 0.1315
-2.750 -0.1814 0.02515 0.01549 -0.0474 0.9536 0.1388
-2.500 -0.1489 0.02460 0.01484 -0.0484 0.9487 0.1482
-2.250 -0.1107 0.02396 0.01431 -0.0506 0.9448 0.1618
-2.000 -0.0797 0.02342 0.01392 -0.0516 0.9400 0.1775
-1.750 -0.0518 0.02298 0.01364 -0.0520 0.9344 0.1973
-1.500 -0.0164 0.02239 0.01332 -0.0537 0.9303 0.2310
-1.250 0.0156 0.02008 0.01318 -0.0550 0.9279 0.6311
-1.000 0.0091 0.02033 0.01407 -0.0449 0.9202 0.8747
-0.750 0.1538 0.02088 0.01435 -0.0635 0.9279 0.9969
-0.500 0.1681 0.02115 0.01453 -0.0622 0.9187 1.0000
-0.250 0.2022 0.02130 0.01458 -0.0641 0.9132 1.0000
0.000 0.2089 0.02164 0.01486 -0.0612 0.9040 1.0000
0.250 0.2348 0.02188 0.01502 -0.0616 0.8974 1.0000
0.500 0.2787 0.02209 0.01516 -0.0650 0.8938 1.0000
0.750 0.2653 0.02260 0.01563 -0.0585 0.8817 1.0000
1.000 0.3060 0.02285 0.01584 -0.0612 0.8775 1.0000
1.250 0.3085 0.02344 0.01638 -0.0576 0.8675 1.0000
1.500 0.3432 0.02375 0.01666 -0.0592 0.8617 1.0000
1.750 0.3888 0.02395 0.01687 -0.0624 0.8581 1.0000
2.000 0.3899 0.02469 0.01758 -0.0586 0.8462 1.0000
2.250 0.4351 0.02478 0.01770 -0.0616 0.8414 1.0000
2.500 0.4581 0.02495 0.01788 -0.0605 0.8279 1.0000
2.750 0.5075 0.02383 0.01680 -0.0617 0.8114 1.0000
3.000 0.5515 0.02245 0.01545 -0.0615 0.7947 1.0000
3.250 0.5883 0.02128 0.01431 -0.0604 0.7801 1.0000
3.500 0.6247 0.02009 0.01313 -0.0593 0.7685 1.0000
3.750 0.6490 0.01966 0.01275 -0.0574 0.7531 1.0000
4.000 0.6752 0.01904 0.01217 -0.0556 0.7373 1.0000
4.250 0.7024 0.01826 0.01143 -0.0536 0.7196 1.0000
4.500 0.7304 0.01732 0.01051 -0.0515 0.6984 1.0000
4.750 0.7535 0.01664 0.00988 -0.0491 0.6666 1.0000
5.000 0.7754 0.01595 0.00911 -0.0461 0.6099 1.0000
5.250 0.7806 0.01657 0.00842 -0.0403 0.3722 1.0000
5.500 0.7808 0.01896 0.00963 -0.0364 0.2438 1.0000
5.750 0.7941 0.02041 0.01065 -0.0342 0.2073 1.0000
6.000 0.8125 0.02158 0.01160 -0.0328 0.1849 1.0000
6.250 0.8333 0.02273 0.01256 -0.0317 0.1687 1.0000
6.500 0.8564 0.02395 0.01359 -0.0310 0.1559 1.0000
6.750 0.8807 0.02497 0.01457 -0.0304 0.1452 1.0000
7.000 0.9070 0.02626 0.01582 -0.0302 0.1368 1.0000
7.250 0.9329 0.02746 0.01691 -0.0300 0.1292 1.0000
7.500 0.9603 0.02894 0.01844 -0.0298 0.1233 1.0000
7.750 0.9862 0.03021 0.01979 -0.0295 0.1179 1.0000
8.000 1.0148 0.03218 0.02159 -0.0300 0.1131 1.0000
8.250 1.0374 0.03355 0.02330 -0.0290 0.1096 1.0000
8.500 1.0611 0.03522 0.02517 -0.0284 0.1063 1.0000
8.750 1.0851 0.03684 0.02683 -0.0281 0.1030 1.0000
9.000 1.1086 0.03944 0.02946 -0.0280 0.1001 1.0000
9.250 1.1250 0.04156 0.03204 -0.0264 0.0986 1.0000
9.500 1.1397 0.04405 0.03495 -0.0248 0.0970 1.0000
9.750 1.1521 0.04664 0.03792 -0.0231 0.0952 1.0000
10.000 1.1633 0.04927 0.04085 -0.0215 0.0935 1.0000
10.250 1.1708 0.05240 0.04432 -0.0196 0.0928 1.0000
10.500 1.1721 0.05609 0.04840 -0.0173 0.0928 1.0000
10.750 1.1669 0.06008 0.05277 -0.0146 0.0929 1.0000
11.000 1.1535 0.06441 0.05748 -0.0116 0.0935 1.0000
11.250 1.1281 0.06902 0.06245 -0.0081 0.0947 1.0000
11.500 1.0983 0.07374 0.06743 -0.0053 0.0960 1.0000
11.750 1.0701 0.07897 0.07287 -0.0040 0.0973 1.0000
12.000 1.0469 0.08458 0.07862 -0.0040 0.0985 1.0000
12.250 1.0374 0.09013 0.08424 -0.0045 0.0996 1.0000
|
Polar data table (+)
Polar graphs
<< Back to RUTAN CANARD AIRFOIL (amsoil1-il)