BOEING-VERTOL VR-5 AIRFOIL (vr5-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: BOEING-VERTOL VR-5 AIRFOIL (vr5-il) Reynolds number: 1,000,000 Max Cl/Cd: 136.12 at α=5.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-vr5-il-1000000.txt Download as CSV file: xf-vr5-il-1000000.csv |
XFOIL Version 6.96
Calculated polar for: BOEING-VERTOL VR-5 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 1.000 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.500 -0.5007 0.09435 0.09243 -0.0030 0.8182 0.0134
-9.250 -0.4973 0.09090 0.08895 -0.0049 0.8105 0.0135
-9.000 -0.4933 0.08749 0.08553 -0.0069 0.8015 0.0137
-8.750 -0.4905 0.08364 0.08166 -0.0096 0.7924 0.0138
-8.500 -0.4892 0.07917 0.07716 -0.0136 0.7821 0.0140
-8.250 -0.4928 0.07207 0.07003 -0.0237 0.7711 0.0141
-8.000 -0.4884 0.06664 0.06450 -0.0292 0.7616 0.0143
-7.750 -0.5706 0.02020 0.01576 -0.0406 0.7616 0.0120
-7.500 -0.5476 0.01793 0.01309 -0.0404 0.7526 0.0124
-7.250 -0.5222 0.01655 0.01146 -0.0403 0.7440 0.0128
-7.000 -0.4958 0.01563 0.01031 -0.0403 0.7357 0.0130
-6.750 -0.4707 0.01403 0.00845 -0.0401 0.7280 0.0135
-6.500 -0.4437 0.01341 0.00772 -0.0401 0.7206 0.0140
-6.250 -0.4156 0.01310 0.00737 -0.0403 0.7138 0.0145
-6.000 -0.3877 0.01276 0.00695 -0.0403 0.7071 0.0152
-5.750 -0.3597 0.01237 0.00648 -0.0404 0.7010 0.0159
-5.500 -0.3313 0.01215 0.00617 -0.0406 0.6946 0.0166
-5.250 -0.3042 0.01149 0.00544 -0.0406 0.6888 0.0177
-5.000 -0.2756 0.01134 0.00528 -0.0408 0.6826 0.0186
-4.750 -0.2472 0.01115 0.00502 -0.0409 0.6764 0.0197
-4.500 -0.2184 0.01099 0.00482 -0.0411 0.6704 0.0208
-4.250 -0.1909 0.01043 0.00421 -0.0412 0.6640 0.0224
-4.000 -0.1621 0.01033 0.00409 -0.0414 0.6582 0.0238
-3.750 -0.1332 0.01022 0.00396 -0.0416 0.6521 0.0254
-3.500 -0.1049 0.00994 0.00362 -0.0418 0.6462 0.0270
-3.250 -0.0766 0.00961 0.00328 -0.0420 0.6403 0.0288
-3.000 -0.0478 0.00944 0.00309 -0.0422 0.6340 0.0304
-2.750 -0.0190 0.00932 0.00292 -0.0424 0.6283 0.0319
-2.500 0.0101 0.00924 0.00283 -0.0427 0.6227 0.0330
-2.250 0.0384 0.00880 0.00233 -0.0429 0.6171 0.0354
-2.000 0.0672 0.00863 0.00213 -0.0432 0.6116 0.0372
-1.750 0.0963 0.00849 0.00197 -0.0434 0.6057 0.0390
-1.500 0.1252 0.00840 0.00182 -0.0437 0.6003 0.0407
-1.250 0.1543 0.00830 0.00170 -0.0440 0.5959 0.0420
-1.000 0.1834 0.00812 0.00152 -0.0442 0.5914 0.0475
-0.750 0.2125 0.00802 0.00141 -0.0445 0.5869 0.0575
-0.500 0.2408 0.00752 0.00133 -0.0450 0.5825 0.2004
-0.250 0.2668 0.00617 0.00122 -0.0456 0.5782 0.6158
0.000 0.2953 0.00605 0.00123 -0.0458 0.5737 0.6700
0.250 0.3239 0.00601 0.00125 -0.0460 0.5692 0.7061
0.500 0.3525 0.00594 0.00129 -0.0461 0.5654 0.7421
0.750 0.3811 0.00592 0.00133 -0.0462 0.5615 0.7724
1.000 0.4096 0.00592 0.00136 -0.0464 0.5576 0.7925
1.250 0.4376 0.00593 0.00141 -0.0464 0.5538 0.8197
1.500 0.4653 0.00588 0.00147 -0.0463 0.5511 0.8469
2.000 0.5212 0.00588 0.00157 -0.0463 0.5442 0.8823
2.250 0.5488 0.00593 0.00163 -0.0462 0.5406 0.8979
2.500 0.5764 0.00597 0.00169 -0.0461 0.5370 0.9131
2.750 0.6038 0.00597 0.00174 -0.0460 0.5330 0.9278
3.000 0.6304 0.00598 0.00178 -0.0457 0.5283 0.9432
3.250 0.6574 0.00603 0.00182 -0.0455 0.5230 0.9606
3.500 0.6902 0.00601 0.00185 -0.0466 0.5185 0.9795
3.750 0.7244 0.00606 0.00190 -0.0481 0.5135 1.0000
4.000 0.7535 0.00616 0.00198 -0.0486 0.5075 1.0000
4.250 0.7829 0.00620 0.00205 -0.0490 0.5024 1.0000
4.500 0.8119 0.00629 0.00213 -0.0495 0.4963 1.0000
4.750 0.8410 0.00637 0.00221 -0.0499 0.4868 1.0000
5.000 0.8697 0.00647 0.00229 -0.0502 0.4745 1.0000
5.250 0.8984 0.00660 0.00239 -0.0506 0.4586 1.0000
5.500 0.9264 0.00682 0.00252 -0.0510 0.4342 1.0000
5.750 0.9538 0.00715 0.00272 -0.0513 0.4000 1.0000
6.000 0.9803 0.00761 0.00301 -0.0515 0.3630 1.0000
6.250 1.0048 0.00839 0.00347 -0.0516 0.2979 1.0000
6.500 1.0298 0.00903 0.00391 -0.0517 0.2585 1.0000
6.750 1.0557 0.00949 0.00426 -0.0519 0.2334 1.0000
7.000 1.0809 0.01004 0.00466 -0.0520 0.2019 1.0000
7.250 1.1017 0.01114 0.00537 -0.0518 0.1335 1.0000
7.500 1.1248 0.01186 0.00594 -0.0517 0.1043 1.0000
7.750 1.1473 0.01261 0.00651 -0.0515 0.0746 1.0000
8.000 1.1704 0.01323 0.00705 -0.0513 0.0591 1.0000
8.250 1.1936 0.01380 0.00757 -0.0512 0.0473 1.0000
8.500 1.2161 0.01440 0.00811 -0.0509 0.0368 1.0000
8.750 1.2380 0.01503 0.00870 -0.0506 0.0288 1.0000
9.000 1.2588 0.01574 0.00937 -0.0501 0.0226 1.0000
9.250 1.2800 0.01635 0.01000 -0.0497 0.0196 1.0000
9.500 1.2992 0.01710 0.01077 -0.0491 0.0175 1.0000
9.750 1.3192 0.01772 0.01144 -0.0485 0.0165 1.0000
10.000 1.3372 0.01842 0.01218 -0.0478 0.0156 1.0000
10.250 1.3511 0.01931 0.01310 -0.0467 0.0147 1.0000
10.500 1.3608 0.02059 0.01445 -0.0454 0.0139 1.0000
10.750 1.3738 0.02164 0.01557 -0.0445 0.0136 1.0000
11.000 1.3868 0.02273 0.01672 -0.0437 0.0133 1.0000
11.250 1.3986 0.02394 0.01800 -0.0429 0.0130 1.0000
11.500 1.4094 0.02527 0.01940 -0.0421 0.0127 1.0000
11.750 1.4192 0.02672 0.02092 -0.0413 0.0124 1.0000
12.000 1.4281 0.02829 0.02255 -0.0406 0.0121 1.0000
12.250 1.4361 0.02996 0.02429 -0.0399 0.0118 1.0000
12.500 1.4427 0.03180 0.02620 -0.0393 0.0115 1.0000
12.750 1.4459 0.03400 0.02846 -0.0386 0.0112 1.0000
13.000 1.4449 0.03665 0.03120 -0.0378 0.0109 1.0000
13.250 1.4393 0.03976 0.03443 -0.0369 0.0106 1.0000
13.500 1.4388 0.04242 0.03719 -0.0363 0.0105 1.0000
13.750 1.4437 0.04456 0.03941 -0.0360 0.0104 1.0000
14.000 1.4469 0.04690 0.04183 -0.0356 0.0103 1.0000
14.250 1.4499 0.04934 0.04436 -0.0354 0.0102 1.0000
14.500 1.4519 0.05195 0.04705 -0.0352 0.0101 1.0000
14.750 1.4534 0.05467 0.04986 -0.0352 0.0099 1.0000
15.000 1.4549 0.05746 0.05274 -0.0352 0.0097 1.0000
15.250 1.4557 0.06039 0.05575 -0.0354 0.0096 1.0000
15.500 1.4567 0.06332 0.05877 -0.0356 0.0094 1.0000
15.750 1.4568 0.06642 0.06195 -0.0359 0.0092 1.0000
16.000 1.4562 0.06970 0.06531 -0.0363 0.0090 1.0000
16.250 1.4546 0.07314 0.06883 -0.0369 0.0089 1.0000
16.500 1.4519 0.07675 0.07254 -0.0375 0.0088 1.0000
16.750 1.4484 0.08054 0.07641 -0.0382 0.0086 1.0000
17.000 1.4441 0.08452 0.08047 -0.0391 0.0085 1.0000
17.250 1.4385 0.08872 0.08476 -0.0400 0.0084 1.0000
17.500 1.4310 0.09321 0.08934 -0.0411 0.0083 1.0000
17.750 1.4215 0.09804 0.09428 -0.0423 0.0082 1.0000
18.000 1.4096 0.10329 0.09964 -0.0437 0.0081 1.0000
|
Polar data table (+)
Polar graphs
<< Back to BOEING-VERTOL VR-5 AIRFOIL (vr5-il)