BOEING-VERTOL VR-15 AIRFOIL (vr15-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
|---|---|
|
Airfoil: BOEING-VERTOL VR-15 AIRFOIL (vr15-il) Reynolds number: 100,000 Max Cl/Cd: 38.24 at α=3.25° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-vr15-il-100000.txt Download as CSV file: xf-vr15-il-100000.csv |
XFOIL Version 6.96
Calculated polar for: BOEING-VERTOL VR-15 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
------ -------- --------- --------- -------- -------- --------
-8.750 -0.5848 0.10093 0.09624 0.0050 1.0000 0.0833
-8.500 -0.5975 0.09635 0.09173 -0.0037 1.0000 0.0848
-8.250 -0.6128 0.09234 0.08762 -0.0111 1.0000 0.0855
-8.000 -0.5952 0.08682 0.08225 -0.0062 1.0000 0.0883
-7.750 -0.5860 0.08345 0.07890 -0.0053 1.0000 0.0924
-7.500 -0.5866 0.07902 0.07441 -0.0096 1.0000 0.0973
-7.250 -0.5958 0.07418 0.06928 -0.0155 1.0000 0.1009
-7.000 -0.5775 0.07030 0.06562 -0.0124 1.0000 0.1056
-6.750 -0.5811 0.06707 0.06188 -0.0164 1.0000 0.1150
-6.500 -0.5621 0.06278 0.05793 -0.0139 1.0000 0.1224
-6.250 -0.5538 0.05910 0.05415 -0.0142 1.0000 0.1342
-6.000 -0.5438 0.05574 0.05070 -0.0140 1.0000 0.1480
-5.750 -0.5324 0.05257 0.04742 -0.0134 1.0000 0.1629
-5.500 -0.5202 0.04967 0.04447 -0.0123 1.0000 0.1802
-5.250 -0.5104 0.04695 0.04166 -0.0110 1.0000 0.2052
-5.000 -0.4996 0.04466 0.03945 -0.0085 1.0000 0.2361
-4.750 -0.4898 0.04253 0.03738 -0.0057 1.0000 0.2689
-4.500 -0.4499 0.03256 0.02492 -0.0063 1.0000 0.1021
-4.250 -0.4299 0.02923 0.02118 -0.0039 1.0000 0.0903
-4.000 -0.4107 0.02708 0.01861 -0.0018 1.0000 0.0897
-3.750 -0.3891 0.02580 0.01671 0.0007 1.0000 0.0861
-3.500 -0.3675 0.02417 0.01478 0.0023 1.0000 0.0858
-3.250 -0.3457 0.02202 0.01259 0.0033 1.0000 0.0896
-3.000 -0.3229 0.02082 0.01130 0.0044 1.0000 0.0925
-2.750 -0.2992 0.01972 0.01008 0.0056 1.0000 0.0945
-2.500 -0.2762 0.01892 0.00917 0.0067 1.0000 0.0987
-2.250 -0.2536 0.01789 0.00823 0.0076 1.0000 0.1045
-2.000 -0.2318 0.01716 0.00755 0.0088 1.0000 0.1099
-1.750 -0.0500 0.01251 0.00602 -0.0179 1.0000 1.0000
-1.500 -0.0406 0.01257 0.00591 -0.0149 1.0000 1.0000
-1.250 -0.0318 0.01268 0.00587 -0.0117 1.0000 1.0000
-1.000 -0.0228 0.01281 0.00588 -0.0087 1.0000 1.0000
-0.750 -0.0129 0.01297 0.00591 -0.0059 1.0000 1.0000
-0.500 -0.0020 0.01315 0.00597 -0.0033 1.0000 1.0000
-0.250 0.0099 0.01335 0.00608 -0.0009 1.0000 1.0000
0.000 0.0227 0.01357 0.00622 0.0013 1.0000 1.0000
0.250 0.0363 0.01382 0.00639 0.0032 1.0000 1.0000
0.500 0.0814 0.01409 0.00660 -0.0009 0.9925 1.0000
0.750 0.1439 0.01429 0.00677 -0.0083 0.9795 1.0000
1.000 0.2085 0.01436 0.00685 -0.0158 0.9657 1.0000
1.250 0.2754 0.01426 0.00681 -0.0235 0.9508 1.0000
1.500 0.3315 0.01404 0.00668 -0.0285 0.9321 1.0000
1.750 0.3708 0.01383 0.00654 -0.0298 0.9092 1.0000
2.000 0.4001 0.01358 0.00633 -0.0286 0.8846 1.0000
2.250 0.4210 0.01335 0.00613 -0.0257 0.8561 1.0000
2.500 0.4392 0.01314 0.00595 -0.0223 0.8221 1.0000
2.750 0.4573 0.01291 0.00568 -0.0187 0.7815 1.0000
3.000 0.4752 0.01277 0.00543 -0.0151 0.7206 1.0000
3.250 0.4921 0.01287 0.00510 -0.0113 0.6004 1.0000
3.500 0.5094 0.01380 0.00513 -0.0086 0.4629 1.0000
3.750 0.5310 0.01472 0.00555 -0.0074 0.3985 1.0000
4.000 0.5538 0.01552 0.00603 -0.0065 0.3629 1.0000
4.250 0.5774 0.01626 0.00658 -0.0057 0.3359 1.0000
4.500 0.6013 0.01700 0.00721 -0.0050 0.3147 1.0000
4.750 0.6256 0.01773 0.00788 -0.0044 0.2974 1.0000
5.000 0.6499 0.01849 0.00860 -0.0037 0.2823 1.0000
5.250 0.6739 0.01921 0.00930 -0.0031 0.2671 1.0000
5.500 0.6976 0.01990 0.01001 -0.0024 0.2521 1.0000
5.750 0.7212 0.02060 0.01073 -0.0017 0.2380 1.0000
6.000 0.7451 0.02141 0.01158 -0.0011 0.2262 1.0000
6.250 0.7689 0.02229 0.01246 -0.0004 0.2147 1.0000
6.500 0.7921 0.02311 0.01331 0.0003 0.2017 1.0000
6.750 0.8144 0.02383 0.01412 0.0011 0.1869 1.0000
7.000 0.8361 0.02457 0.01498 0.0021 0.1714 1.0000
7.250 0.8573 0.02534 0.01588 0.0031 0.1555 1.0000
7.500 0.8780 0.02615 0.01683 0.0041 0.1401 1.0000
7.750 0.8987 0.02723 0.01802 0.0052 0.1265 1.0000
8.000 0.9187 0.02826 0.01913 0.0063 0.1135 1.0000
8.250 0.9389 0.02962 0.02049 0.0074 0.1031 1.0000
8.500 0.9581 0.03100 0.02212 0.0087 0.0939 1.0000
8.750 0.9754 0.03293 0.02434 0.0100 0.0862 1.0000
9.000 0.9940 0.03454 0.02609 0.0111 0.0806 1.0000
9.250 1.0080 0.03762 0.02951 0.0126 0.0771 1.0000
9.500 1.0169 0.04071 0.03318 0.0146 0.0735 1.0000
9.750 1.0317 0.04240 0.03499 0.0158 0.0693 1.0000
10.000 1.0432 0.04569 0.03833 0.0168 0.0667 1.0000
10.250 1.0398 0.04992 0.04313 0.0191 0.0660 1.0000
10.500 1.0305 0.05456 0.04826 0.0213 0.0657 1.0000
10.750 1.0157 0.05940 0.05348 0.0231 0.0657 1.0000
11.000 0.9962 0.06417 0.05851 0.0246 0.0659 1.0000
11.250 0.9735 0.06880 0.06330 0.0257 0.0662 1.0000
11.500 0.9515 0.07422 0.06883 0.0249 0.0665 1.0000
11.750 0.7913 0.11628 0.11093 -0.0068 0.0928 1.0000
12.000 0.7956 0.12038 0.11505 -0.0068 0.0920 1.0000
|
Polar data table (+)
Polar graphs
<< Back to BOEING-VERTOL VR-15 AIRFOIL (vr15-il)