BOEING-VERTOL VR-7 AIRFOIL (vr7-il) Xfoil prediction polar at RE=50,000 Ncrit=5
| Details | Polar file |
|---|---|
|
Airfoil: BOEING-VERTOL VR-7 AIRFOIL (vr7-il) Reynolds number: 50,000 Max Cl/Cd: 32.11 at α=8.5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-vr7-il-50000-n5.txt Download as CSV file: xf-vr7-il-50000-n5.csv |
XFOIL Version 6.96
Calculated polar for: BOEING-VERTOL VR-7 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 = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-9.750 -0.3788 0.10875 0.10187 -0.0255 1.0000 0.1427
-9.500 -0.4025 0.10695 0.10024 -0.0301 1.0000 0.1492
-9.250 -0.4120 0.09277 0.08602 -0.0418 1.0000 0.0683
-9.000 -0.4008 0.08963 0.08292 -0.0405 1.0000 0.0673
-8.750 -0.3982 0.08596 0.07932 -0.0411 1.0000 0.0659
-8.500 -0.4023 0.08181 0.07526 -0.0431 1.0000 0.0642
-8.000 -0.4495 0.07136 0.06472 -0.0483 1.0000 0.0575
-7.750 -0.4499 0.06829 0.06168 -0.0473 1.0000 0.0571
-7.500 -0.4538 0.06529 0.05869 -0.0459 1.0000 0.0567
-7.250 -0.4594 0.06257 0.05596 -0.0440 1.0000 0.0562
-7.000 -0.4676 0.06030 0.05366 -0.0413 1.0000 0.0556
-6.750 -0.4693 0.05755 0.05078 -0.0400 0.9960 0.0550
-6.500 -0.4450 0.05282 0.04564 -0.0439 0.9804 0.0540
-6.250 -0.4197 0.04857 0.04089 -0.0468 0.9660 0.0533
-6.000 -0.3921 0.04476 0.03655 -0.0492 0.9532 0.0532
-5.750 -0.3612 0.04148 0.03268 -0.0515 0.9421 0.0547
-5.500 -0.3319 0.03875 0.02936 -0.0528 0.9300 0.0559
-5.250 -0.3002 0.03627 0.02631 -0.0539 0.9192 0.0568
-5.000 -0.2639 0.03398 0.02349 -0.0555 0.9108 0.0574
-4.750 -0.2334 0.03205 0.02134 -0.0561 0.8995 0.0586
-4.500 -0.2006 0.03050 0.01966 -0.0571 0.8898 0.0616
-4.250 -0.1659 0.02914 0.01803 -0.0580 0.8809 0.0657
-4.000 -0.1343 0.02802 0.01666 -0.0581 0.8708 0.0684
-3.750 -0.1002 0.02673 0.01535 -0.0589 0.8631 0.0728
-3.500 -0.0735 0.02601 0.01440 -0.0585 0.8522 0.0795
-3.250 -0.0460 0.02509 0.01347 -0.0585 0.8431 0.0869
-3.000 -0.0188 0.02430 0.01262 -0.0584 0.8339 0.0993
-2.750 0.0046 0.02352 0.01192 -0.0577 0.8241 0.1234
-2.500 0.0204 0.02129 0.01132 -0.0563 0.8161 0.3944
-2.250 0.0343 0.02048 0.01189 -0.0507 0.8070 0.7567
-2.000 0.1596 0.02083 0.01189 -0.0630 0.8046 0.9567
-1.750 0.2099 0.02062 0.01131 -0.0676 0.7951 0.9753
-1.500 0.2564 0.02040 0.01075 -0.0718 0.7853 0.9928
-1.250 0.2834 0.02037 0.01051 -0.0726 0.7744 1.0000
-1.000 0.3002 0.02042 0.01037 -0.0712 0.7648 1.0000
-0.750 0.3165 0.02055 0.01032 -0.0696 0.7548 1.0000
-0.500 0.3335 0.02074 0.01036 -0.0682 0.7452 1.0000
-0.250 0.3540 0.02089 0.01033 -0.0671 0.7370 1.0000
0.000 0.3717 0.02119 0.01051 -0.0658 0.7272 1.0000
0.250 0.3944 0.02134 0.01049 -0.0649 0.7198 1.0000
0.500 0.4119 0.02173 0.01079 -0.0635 0.7098 1.0000
0.750 0.4340 0.02197 0.01090 -0.0625 0.7022 1.0000
1.000 0.4534 0.02233 0.01118 -0.0613 0.6928 1.0000
1.250 0.4748 0.02264 0.01141 -0.0602 0.6848 1.0000
1.500 0.4956 0.02299 0.01168 -0.0591 0.6762 1.0000
1.750 0.5163 0.02337 0.01201 -0.0579 0.6677 1.0000
2.000 0.5383 0.02369 0.01227 -0.0569 0.6596 1.0000
2.250 0.5582 0.02414 0.01270 -0.0557 0.6508 1.0000
2.500 0.5814 0.02445 0.01295 -0.0548 0.6431 1.0000
2.750 0.6005 0.02498 0.01351 -0.0536 0.6340 1.0000
3.000 0.6246 0.02526 0.01375 -0.0528 0.6267 1.0000
3.250 0.6427 0.02587 0.01440 -0.0515 0.6172 1.0000
3.500 0.6678 0.02612 0.01464 -0.0508 0.6103 1.0000
3.750 0.6849 0.02681 0.01540 -0.0495 0.6004 1.0000
4.000 0.7111 0.02701 0.01559 -0.0488 0.5939 1.0000
4.250 0.7271 0.02778 0.01645 -0.0474 0.5836 1.0000
4.500 0.7534 0.02797 0.01668 -0.0468 0.5772 1.0000
4.750 0.7695 0.02874 0.01755 -0.0455 0.5669 1.0000
5.000 0.7911 0.02921 0.01810 -0.0445 0.5588 1.0000
5.250 0.8127 0.02964 0.01863 -0.0435 0.5501 1.0000
5.500 0.8306 0.03025 0.01934 -0.0421 0.5398 1.0000
5.750 0.8599 0.03008 0.01921 -0.0415 0.5321 1.0000
6.000 0.8753 0.03073 0.01999 -0.0399 0.5200 1.0000
6.250 0.8938 0.03122 0.02061 -0.0385 0.5090 1.0000
6.500 0.9234 0.03103 0.02050 -0.0380 0.5014 1.0000
6.750 0.9368 0.03191 0.02154 -0.0363 0.4897 1.0000
7.000 0.9546 0.03242 0.02219 -0.0348 0.4784 1.0000
7.250 0.9777 0.03249 0.02240 -0.0336 0.4672 1.0000
7.500 1.0036 0.03233 0.02234 -0.0326 0.4557 1.0000
7.750 1.0177 0.03298 0.02316 -0.0308 0.4424 1.0000
8.000 1.0330 0.03352 0.02389 -0.0290 0.4291 1.0000
8.250 1.0506 0.03366 0.02414 -0.0272 0.4136 1.0000
8.500 1.0710 0.03335 0.02386 -0.0253 0.3954 1.0000
8.750 1.0783 0.03394 0.02455 -0.0226 0.3757 1.0000
9.000 1.0861 0.03454 0.02526 -0.0200 0.3558 1.0000
9.250 1.0933 0.03518 0.02594 -0.0173 0.3365 1.0000
9.500 1.0967 0.03610 0.02692 -0.0145 0.3178 1.0000
9.750 1.0976 0.03738 0.02827 -0.0120 0.2984 1.0000
10.000 1.1002 0.03866 0.02957 -0.0098 0.2787 1.0000
10.250 1.1015 0.04019 0.03107 -0.0079 0.2587 1.0000
10.500 1.0999 0.04216 0.03304 -0.0063 0.2379 1.0000
10.750 1.0981 0.04425 0.03502 -0.0049 0.2181 1.0000
11.000 1.0951 0.04670 0.03740 -0.0039 0.1996 1.0000
11.250 1.0923 0.04928 0.03989 -0.0030 0.1832 1.0000
11.500 1.0897 0.05200 0.04255 -0.0023 0.1684 1.0000
11.750 1.0877 0.05479 0.04532 -0.0018 0.1551 1.0000
12.000 1.0858 0.05766 0.04817 -0.0014 0.1432 1.0000
12.250 1.0851 0.06042 0.05088 -0.0011 0.1333 1.0000
12.500 1.0841 0.06346 0.05401 -0.0010 0.1235 1.0000
12.750 1.0842 0.06638 0.05697 -0.0009 0.1154 1.0000
13.000 1.0845 0.06931 0.05996 -0.0009 0.1079 1.0000
13.250 1.0846 0.07249 0.06325 -0.0010 0.1013 1.0000
13.500 1.0858 0.07541 0.06620 -0.0012 0.0954 1.0000
13.750 1.0863 0.07870 0.06963 -0.0014 0.0903 1.0000
14.000 1.0854 0.08225 0.07335 -0.0020 0.0857 1.0000
14.250 1.0940 0.08423 0.07519 -0.0015 0.0812 1.0000
14.500 1.0864 0.08923 0.08056 -0.0029 0.0784 1.0000
14.750 1.0789 0.09428 0.08589 -0.0045 0.0758 1.0000
15.000 1.0723 0.09918 0.09099 -0.0061 0.0734 1.0000
15.250 1.0787 0.10163 0.09341 -0.0063 0.0701 1.0000
15.500 1.0686 0.10742 0.09938 -0.0086 0.0687 1.0000
15.750 1.0454 0.11606 0.10830 -0.0131 0.0681 1.0000
16.000 1.0168 0.12659 0.11903 -0.0190 0.0681 1.0000
16.250 0.9795 0.14036 0.13295 -0.0269 0.0682 1.0000
|
Polar data table (+)
Polar graphs
<< Back to BOEING-VERTOL VR-7 AIRFOIL (vr7-il)