BOEING-VERTOL VR-11X AIRFOIL (vr11x-il) Xfoil prediction polar at RE=500,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: BOEING-VERTOL VR-11X AIRFOIL (vr11x-il) Reynolds number: 500,000 Max Cl/Cd: 73.21 at α=8.5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-vr11x-il-500000-n5.txt Download as CSV file: xf-vr11x-il-500000-n5.csv |
XFOIL Version 6.96
Calculated polar for: BOEING-VERTOL VR-11X 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 = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.250 -0.4973 0.06488 0.06217 -0.0428 0.8687 0.0105
-8.000 -0.4968 0.06085 0.05804 -0.0438 0.8678 0.0103
-7.750 -0.4942 0.05669 0.05375 -0.0443 0.8669 0.0102
-7.500 -0.4903 0.05215 0.04904 -0.0443 0.8661 0.0101
-7.250 -0.4849 0.04711 0.04375 -0.0437 0.8652 0.0102
-7.000 -0.4769 0.04242 0.03880 -0.0424 0.8644 0.0103
-6.750 -0.4653 0.03888 0.03502 -0.0413 0.8638 0.0102
-6.500 -0.4519 0.03544 0.03133 -0.0400 0.8631 0.0099
-6.250 -0.4377 0.03172 0.02728 -0.0383 0.8622 0.0096
-6.000 -0.4233 0.02754 0.02267 -0.0361 0.8611 0.0093
-5.750 -0.4072 0.02331 0.01791 -0.0337 0.8603 0.0090
-5.500 -0.3865 0.02046 0.01463 -0.0322 0.8594 0.0090
-5.250 -0.3626 0.01874 0.01263 -0.0313 0.8584 0.0092
-5.000 -0.3376 0.01745 0.01112 -0.0306 0.8573 0.0093
-4.750 -0.3119 0.01640 0.00990 -0.0301 0.8563 0.0096
-4.500 -0.2859 0.01553 0.00889 -0.0296 0.8555 0.0099
-4.250 -0.2597 0.01482 0.00807 -0.0292 0.8549 0.0102
-4.000 -0.2333 0.01425 0.00741 -0.0289 0.8543 0.0105
-3.750 -0.2082 0.01368 0.00677 -0.0282 0.8527 0.0108
-3.500 -0.1844 0.01315 0.00619 -0.0272 0.8499 0.0111
-3.250 -0.1578 0.01273 0.00578 -0.0270 0.8464 0.0115
-3.000 -0.1313 0.01241 0.00545 -0.0266 0.8433 0.0122
-2.750 -0.1053 0.01209 0.00509 -0.0261 0.8403 0.0131
-2.500 -0.0799 0.01177 0.00471 -0.0254 0.8372 0.0136
-2.250 -0.0540 0.01135 0.00429 -0.0249 0.8331 0.0142
-2.000 -0.0274 0.01102 0.00396 -0.0245 0.8277 0.0151
-1.750 -0.0016 0.01072 0.00362 -0.0238 0.8227 0.0163
-1.500 0.0259 0.01044 0.00334 -0.0236 0.8166 0.0178
-1.250 0.0530 0.01017 0.00307 -0.0233 0.8101 0.0206
-1.000 0.0803 0.00991 0.00285 -0.0230 0.8039 0.0275
-0.750 0.1072 0.00959 0.00263 -0.0228 0.7945 0.0556
-0.500 0.1299 0.00874 0.00242 -0.0220 0.7829 0.2441
-0.250 0.1451 0.00745 0.00223 -0.0198 0.7697 0.5690
0.000 0.1586 0.00629 0.00207 -0.0162 0.7403 0.8411
0.500 0.2889 0.00795 0.00272 -0.0323 0.4953 0.9756
0.750 0.3230 0.00844 0.00284 -0.0341 0.4255 0.9785
1.000 0.3547 0.00887 0.00296 -0.0353 0.3666 0.9819
1.500 0.4177 0.00954 0.00317 -0.0375 0.2900 0.9877
1.750 0.4504 0.00980 0.00326 -0.0388 0.2582 0.9899
2.000 0.4824 0.01008 0.00336 -0.0400 0.2289 0.9924
2.500 0.5470 0.01052 0.00357 -0.0425 0.1984 0.9961
2.750 0.5800 0.01069 0.00367 -0.0438 0.1904 0.9978
3.000 0.6130 0.01087 0.00378 -0.0452 0.1840 0.9993
3.250 0.6421 0.01106 0.00391 -0.0457 0.1784 1.0000
3.500 0.6666 0.01118 0.00403 -0.0452 0.1749 1.0000
3.750 0.6908 0.01134 0.00415 -0.0446 0.1714 1.0000
4.000 0.7148 0.01151 0.00429 -0.0440 0.1684 1.0000
4.250 0.7385 0.01170 0.00445 -0.0433 0.1654 1.0000
4.500 0.7620 0.01191 0.00465 -0.0426 0.1627 1.0000
4.750 0.7858 0.01206 0.00481 -0.0420 0.1611 1.0000
5.000 0.8094 0.01222 0.00498 -0.0413 0.1583 1.0000
5.250 0.8327 0.01240 0.00515 -0.0405 0.1545 1.0000
5.500 0.8555 0.01263 0.00535 -0.0397 0.1501 1.0000
5.750 0.8781 0.01288 0.00557 -0.0389 0.1470 1.0000
6.000 0.9014 0.01303 0.00576 -0.0381 0.1448 1.0000
6.250 0.9245 0.01321 0.00595 -0.0374 0.1415 1.0000
6.500 0.9471 0.01342 0.00616 -0.0365 0.1385 1.0000
6.750 0.9693 0.01365 0.00639 -0.0356 0.1357 1.0000
7.000 0.9909 0.01393 0.00666 -0.0346 0.1328 1.0000
7.250 1.0135 0.01410 0.00689 -0.0338 0.1302 1.0000
7.500 1.0357 0.01431 0.00712 -0.0329 0.1267 1.0000
7.750 1.0574 0.01455 0.00736 -0.0319 0.1230 1.0000
8.000 1.0783 0.01483 0.00763 -0.0308 0.1190 1.0000
8.250 1.0999 0.01505 0.00791 -0.0298 0.1161 1.0000
8.500 1.1209 0.01531 0.00818 -0.0287 0.1119 1.0000
8.750 1.1411 0.01561 0.00847 -0.0276 0.1069 1.0000
9.000 1.1614 0.01589 0.00879 -0.0264 0.1018 1.0000
9.250 1.1808 0.01624 0.00912 -0.0251 0.0951 1.0000
9.500 1.1996 0.01662 0.00948 -0.0237 0.0871 1.0000
9.750 1.2175 0.01705 0.00990 -0.0223 0.0800 1.0000
10.000 1.2346 0.01753 0.01036 -0.0207 0.0733 1.0000
10.250 1.2501 0.01797 0.01082 -0.0188 0.0683 1.0000
10.500 1.2635 0.01849 0.01133 -0.0165 0.0631 1.0000
10.750 1.2778 0.01895 0.01183 -0.0144 0.0593 1.0000
11.000 1.2916 0.01950 0.01240 -0.0124 0.0548 1.0000
11.250 1.3056 0.02008 0.01300 -0.0104 0.0498 1.0000
11.500 1.3187 0.02077 0.01369 -0.0084 0.0440 1.0000
11.750 1.3318 0.02148 0.01440 -0.0065 0.0370 1.0000
12.000 1.3427 0.02237 0.01528 -0.0044 0.0293 1.0000
12.500 1.3661 0.02404 0.01702 -0.0005 0.0241 1.0000
12.750 1.3779 0.02489 0.01793 0.0013 0.0231 1.0000
13.000 1.3889 0.02580 0.01891 0.0031 0.0223 1.0000
13.250 1.3991 0.02678 0.01997 0.0049 0.0215 1.0000
13.500 1.4082 0.02784 0.02112 0.0068 0.0208 1.0000
13.750 1.4175 0.02890 0.02227 0.0085 0.0204 1.0000
14.000 1.4264 0.03001 0.02348 0.0102 0.0201 1.0000
14.250 1.4340 0.03123 0.02480 0.0119 0.0197 1.0000
14.500 1.4404 0.03257 0.02625 0.0136 0.0194 1.0000
14.750 1.4456 0.03403 0.02783 0.0153 0.0191 1.0000
15.000 1.4492 0.03566 0.02956 0.0169 0.0188 1.0000
15.250 1.4511 0.03750 0.03152 0.0184 0.0184 1.0000
15.500 1.4508 0.03960 0.03374 0.0198 0.0181 1.0000
15.750 1.4483 0.04203 0.03628 0.0210 0.0177 1.0000
16.000 1.4436 0.04483 0.03921 0.0218 0.0175 1.0000
16.250 1.4364 0.04815 0.04266 0.0222 0.0172 1.0000
16.500 1.4267 0.05215 0.04680 0.0219 0.0170 1.0000
16.750 1.4140 0.05697 0.05178 0.0207 0.0168 1.0000
17.000 1.3979 0.06276 0.05774 0.0186 0.0167 1.0000
17.250 1.3781 0.06961 0.06476 0.0156 0.0166 1.0000
17.500 1.3543 0.07751 0.07285 0.0119 0.0166 1.0000
17.750 1.3263 0.08636 0.08188 0.0077 0.0166 1.0000
18.000 1.2962 0.09562 0.09132 0.0034 0.0167 1.0000
18.250 1.2654 0.10485 0.10070 -0.0007 0.0169 1.0000
|
Polar data table (+)
Polar graphs
<< Back to BOEING-VERTOL VR-11X AIRFOIL (vr11x-il)