BOEING-VERTOL VR-8 AIRFOIL (vr8-il) Xfoil prediction polar at RE=100,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: BOEING-VERTOL VR-8 AIRFOIL (vr8-il) Reynolds number: 100,000 Max Cl/Cd: 38.68 at α=3.5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-vr8-il-100000-n5.txt Download as CSV file: xf-vr8-il-100000-n5.csv |
XFOIL Version 6.96
Calculated polar for: BOEING-VERTOL VR-8 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 = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.750 -0.5550 0.09312 0.08837 -0.0108 1.0000 0.0576
-8.500 -0.5653 0.08816 0.08349 -0.0178 1.0000 0.0582
-8.000 -0.5836 0.08022 0.07536 -0.0255 1.0000 0.0588
-7.000 -0.5559 0.05710 0.05195 -0.0276 1.0000 0.0328
-6.750 -0.5516 0.05044 0.04479 -0.0279 1.0000 0.0280
-6.500 -0.5407 0.04668 0.04085 -0.0275 1.0000 0.0272
-6.250 -0.5285 0.04262 0.03647 -0.0268 1.0000 0.0262
-6.000 -0.5148 0.03841 0.03180 -0.0256 1.0000 0.0252
-5.750 -0.4988 0.03455 0.02740 -0.0242 1.0000 0.0248
-5.500 -0.4802 0.03222 0.02473 -0.0232 1.0000 0.0268
-5.250 -0.4602 0.02951 0.02137 -0.0216 1.0000 0.0283
-5.000 -0.4390 0.02701 0.01843 -0.0203 1.0000 0.0285
-4.750 -0.4173 0.02480 0.01600 -0.0192 1.0000 0.0291
-4.500 -0.3953 0.02319 0.01423 -0.0182 1.0000 0.0304
-4.250 -0.3726 0.02214 0.01299 -0.0172 1.0000 0.0338
-4.000 -0.3494 0.02083 0.01146 -0.0160 1.0000 0.0350
-3.750 -0.3267 0.01967 0.01019 -0.0148 1.0000 0.0357
-3.500 -0.3060 0.01828 0.00882 -0.0136 1.0000 0.0369
-3.250 -0.2858 0.01743 0.00802 -0.0125 1.0000 0.0401
-3.000 -0.2658 0.01674 0.00728 -0.0112 1.0000 0.0424
-2.750 -0.2303 0.01600 0.00641 -0.0131 0.9888 0.0439
-2.500 -0.1924 0.01542 0.00568 -0.0153 0.9749 0.0459
-2.250 -0.1552 0.01486 0.00505 -0.0174 0.9595 0.0500
-2.000 -0.1182 0.01439 0.00453 -0.0193 0.9430 0.0607
-1.750 -0.0895 0.01199 0.00420 -0.0207 0.9239 0.5335
-1.500 -0.0559 0.01136 0.00403 -0.0206 0.8823 0.6942
-1.250 -0.0247 0.01109 0.00394 -0.0198 0.8474 0.7980
-1.000 0.0228 0.01105 0.00393 -0.0222 0.8246 0.9056
-0.750 0.0617 0.01111 0.00371 -0.0238 0.7913 0.9333
-0.500 0.0979 0.01119 0.00352 -0.0252 0.7592 0.9519
-0.250 0.1286 0.01127 0.00342 -0.0258 0.7418 0.9757
0.000 0.1685 0.01128 0.00323 -0.0283 0.7195 0.9888
0.250 0.2063 0.01129 0.00305 -0.0305 0.6959 1.0000
0.500 0.2297 0.01136 0.00302 -0.0297 0.6813 1.0000
0.750 0.2537 0.01144 0.00301 -0.0291 0.6688 1.0000
1.000 0.2779 0.01154 0.00303 -0.0284 0.6565 1.0000
1.250 0.3022 0.01164 0.00309 -0.0278 0.6451 1.0000
1.500 0.3266 0.01174 0.00317 -0.0272 0.6342 1.0000
1.750 0.3508 0.01187 0.00328 -0.0265 0.6223 1.0000
2.000 0.3745 0.01200 0.00337 -0.0256 0.6039 1.0000
2.250 0.3978 0.01215 0.00345 -0.0247 0.5822 1.0000
2.500 0.4211 0.01230 0.00355 -0.0238 0.5558 1.0000
2.750 0.4445 0.01247 0.00370 -0.0229 0.5286 1.0000
3.000 0.4679 0.01266 0.00385 -0.0220 0.4986 1.0000
3.250 0.4916 0.01286 0.00403 -0.0211 0.4670 1.0000
3.500 0.5129 0.01326 0.00414 -0.0199 0.3903 1.0000
3.750 0.5303 0.01430 0.00445 -0.0185 0.2524 1.0000
4.000 0.5515 0.01504 0.00499 -0.0176 0.1835 1.0000
4.250 0.5704 0.01614 0.00562 -0.0166 0.0884 1.0000
4.500 0.5910 0.01700 0.00626 -0.0157 0.0518 1.0000
4.750 0.6136 0.01759 0.00688 -0.0149 0.0470 1.0000
5.000 0.6360 0.01821 0.00759 -0.0140 0.0445 1.0000
5.250 0.6585 0.01883 0.00841 -0.0131 0.0430 1.0000
5.500 0.6804 0.01952 0.00927 -0.0122 0.0419 1.0000
5.750 0.7017 0.02031 0.01022 -0.0112 0.0411 1.0000
6.000 0.7222 0.02120 0.01126 -0.0101 0.0403 1.0000
6.250 0.7422 0.02218 0.01235 -0.0090 0.0386 1.0000
6.500 0.7612 0.02340 0.01360 -0.0079 0.0365 1.0000
6.750 0.7805 0.02496 0.01523 -0.0067 0.0356 1.0000
7.000 0.8028 0.02623 0.01670 -0.0058 0.0344 1.0000
7.250 0.8247 0.02739 0.01808 -0.0049 0.0311 1.0000
7.500 0.8458 0.02907 0.01982 -0.0042 0.0295 1.0000
7.750 0.8668 0.03067 0.02168 -0.0032 0.0267 1.0000
8.000 0.8865 0.03258 0.02372 -0.0025 0.0250 1.0000
8.250 0.9056 0.03550 0.02704 -0.0014 0.0241 1.0000
8.500 0.9219 0.03860 0.03054 -0.0001 0.0237 1.0000
8.750 0.9379 0.04167 0.03369 0.0005 0.0228 1.0000
9.000 0.9406 0.04539 0.03819 0.0027 0.0217 1.0000
9.250 0.9527 0.04855 0.04132 0.0031 0.0203 1.0000
9.500 0.9433 0.05250 0.04604 0.0056 0.0198 1.0000
9.750 0.9362 0.05641 0.05037 0.0072 0.0198 1.0000
10.000 0.9267 0.06004 0.05428 0.0087 0.0199 1.0000
10.250 0.9107 0.06362 0.05809 0.0100 0.0199 1.0000
10.500 0.8962 0.06747 0.06211 0.0098 0.0200 1.0000
10.750 0.8812 0.07192 0.06670 0.0081 0.0201 1.0000
11.000 0.8701 0.07657 0.07145 0.0058 0.0202 1.0000
11.250 0.8307 0.08692 0.08211 -0.0031 0.0214 1.0000
11.500 0.8123 0.09595 0.09121 -0.0097 0.0220 1.0000
11.750 0.7996 0.10491 0.10016 -0.0154 0.0223 1.0000
12.000 0.7941 0.11208 0.10731 -0.0191 0.0226 1.0000
|
Polar data table (+)
Polar graphs
<< Back to BOEING-VERTOL VR-8 AIRFOIL (vr8-il)